Translational inerter assembly and method for damping movement of a flight control surface

ABSTRACT

There is provided a translational inerter assembly for damping movement of a flight control surface of an aircraft with a support structure. The translational inerter assembly includes a press fit element rotatably disposed within the flight control surface. The translational inerter assembly further includes an inertia element disposed in the press fit element. The translational inerter assembly further includes a torsion bar coupled to the inertia element and to the support structure of the aircraft, such that when the flight control surface rotates, the inertia element translates, and movement of the flight control surface is dampened.

CROSS-REFERENCE TO RELATED APPLICATIONS

The present application is a continuation application of and claimspriority to pending U.S. application Ser. No. 15/867,940, filed on Jan.11, 2018, now U.S. Pat. No. 10,145,434, issued Dec. 4, 2018, andentitled “Translational Inerter Assembly and Method for Damping Movementof a Flight Control Surface”, the entire contents of which is expresslyincorporated by reference herein, and the present continuationapplication is related to contemporaneously filed continuationapplication U.S. patent application Ser. No. 16/159,636, titled “DualRack and Pinion Rotational Inerter System and Method for DampingMovement of a Flight Control Surface of an Aircraft”, filed on Oct. 13,2018, now U.S. Pat. No. 10,352,389, issued Jul. 16, 2019, the entirecontents of which is expressly incorporated by reference herein; andwhich U.S. application Ser. No. 15/867,940 is a continuation-in-partapplication of and claims priority to U.S. application Ser. No.15/159,706, filed on May 19, 2016, now U.S. Pat. No. 10,088,006, issuedOct. 2, 2018, entitled “Rotational Inerter and Method for Damping anActuator”, the entire contents of which is expressly incorporated byreference herein, and which U.S. application Ser. No. 15/867,940 isrelated to continuation-in-part U.S. patent application Ser. No.15/867,988, titled “Dual Rack and Pinion Rotational Inerter System andMethod for Damping Movement of a Flight Control Surface of an Aircraft”,filed on Jan. 11, 2018, now U.S. Pat. No. 10,107,347, issued Oct. 23,2018, the entire contents of which is expressly incorporated byreference herein.

FIELD

The present disclosure relates to actuators and, more particularly, to atranslational inerter assembly and method for damping movement of aflight control surface.

BACKGROUND

Aircraft typically include a flight control system for directional andattitude control of the aircraft in response to commands from a flightcrew or an autopilot. A flight control system may include a plurality ofmovable flight control surfaces such as ailerons on the wings for rollcontrol, elevators on the horizontal tail of the empennage for pitchcontrol, a rudder on the vertical tail of the empennage for yaw control,and other movable control surfaces. Movement of a flight control surfaceis typically effected by one or more actuators mechanically coupledbetween a support structure (e.g., a wing spar) and the flight controlsurface (e.g., an aileron). In many aircraft, the actuators for flightcontrol surfaces are linear hydraulic actuators driven by one or morehydraulic systems which typically operate at a fixed working pressure.

One of the challenges facing aircraft designers is preventing theoccurrence of flutter of the flight control surfaces during flight.Control surface flutter may be described as unstableaerodynamically-induced oscillations of the flight control surface, andmay occur in flight control systems where the operating bandwidth of theflight control system overlaps the resonant frequency of the flightcontrol surface. Unless damped, the oscillations may rapidly increase inamplitude with the potential for undesirable results, includingexceeding the strength capability of the mounting system of the flightcontrol surface and the actuator. Contributing to the potential forcontrol surface flutter is elasticity in the flight control system. Forexample, hydraulic actuators may exhibit a linear spring response underload due to compressibility of the hydraulic fluid. The compressibilityof the hydraulic fluid may be characterized by the cross-sectional areaof the actuator piston, the volume of the hydraulic fluid, and theeffective bulk modulus of elasticity of the hydraulic fluid.

One method of addressing control surface flutter involves designing theflight control system such that the operating bandwidth does not overlapthe resonant frequency of the flight control surface, and may includelimiting the inertia of the load on the actuator and/or increasing thepiston cross-sectional area as a means to react the inertia load.Unfortunately, the above known methods result in an actuator system thatis sized not to provide the actuator with static load-carryingcapability, but rather to provide the actuator with the ability to reactlarger inertia as a means to avoid resonance in the operating bandwidth.As may be appreciated, limiting control surface inertia corresponds to adecrease in control surface area. A decrease in the surface area ofhigher inertia control surfaces of an aircraft empennage may reduce theattitude controllability of the aircraft. As may be appreciated, anincrease in the piston cross-sectional area of an actuator correspondsto an increase in the size and weight of the hydraulic system componentsincluding the size and weight of the actuators, tubing, reservoirs, andother components. The increased size of the actuators may protrudefurther outside of the outer mold line of the aerodynamic surfacesresulting in an increase in aerodynamic drag of an aircraft. The reducedattitude controllability, increased weight of the hydraulic system, andincreased aerodynamic drag may reduce safety, fuel efficiency, range,and/or payload capacity of the aircraft.

As can be seen, there exists a need in the art for a system and methodfor allowing the operating bandwidth of an actuator to match orencompass the resonant frequency of a movable device without oscillatoryresponse.

In addition, flutter suppression is a known challenge for high-pressure,hydraulic, flight-control actuation. High pressure hydraulics systemsface an upper limit due to aero-servo-elasticity which drives a lowerlimit on actuator linear stiffness. That lower limit depends on thekinematics and inertia of the flight control surface.

Known flight control systems and method for addressing fluttersuppression are primarily focused on increasing linear stiffness byincreasing actuator piston diameter, which may cause increased flightcontrol system and aircraft size, weight, and power. Increased flightcontrol system and aircraft size, weight, and power may result inincreased flight fuel costs. Other known flight control systems andmethods for addressing flutter suppression attempt to enhance the activecontrol system performance by increasing the servo bandwidth to operatein the high dynamic resonant frequency range of the actuator and valve.However, such known flight control systems and methods involve the usedof active control elements, such as the actuator and valve size ordiameter, rather than a passive means to change the dynamics of theflight control system. The use of such active control elements mayoverly complicate the control elements and be less space efficient.

As can be seen, there exists a need in the art for an assembly andmethod to address flutter suppression and flutter critical controlsurface applications on aircraft, to dampen movement of flight controlsurfaces, and to optimize a flight control system design in terms ofspace efficiency and changing the dynamic characteristics of thehardware under control rather than complicating the flight controlsystem elements themselves.

SUMMARY

The above-noted needs associated with actuators are specificallyaddressed and alleviated by the present disclosure which provides atranslational inerter assembly for damping movement of a flight controlsurface of an aircraft. The translational inerter assembly comprises apress fit element fixedly disposed within a first end of the flightcontrol surface and rotatably movable with the flight control surface.The translational inerter assembly further comprises an inertia elementcoupled to and installed in the press fit element. The inertia elementhas a plurality of exterior helical splines corresponding to a pluralityof interior helical splines of the press fit element.

The translational inerter assembly further comprises a torsion barhaving a torsion bar first end coupled to and installed in the inertiaelement, and having a torsion bar second end fixedly attached to asupport structure of the aircraft. The torsion bar further has aplurality of exterior linear splines corresponding to a plurality ofinterior linear splines of the inertia element. Rotation of the flightcontrol surface causes translational movement of the inertia element,via the press fit element, along a hinge axis of the flight controlsurface and along the torsion bar, resulting in the translationalinerter assembly damping movement of the flight control surface.

Also disclosed is an aircraft comprising a flight control surfacepivotably coupled to a support structure. The aircraft further comprisesone or more actuators configured to actuate the flight control surface.The aircraft further discloses at least one translational inerterassembly for damping movement of the flight control surface. The atleast one translational inerter assembly is attached to a first end ofthe flight control surface.

The at least one translational inerter assembly comprises a press fitelement fixedly disposed within the first end of the flight controlsurface and rotatably movable with the flight control surface. The atleast one translational inerter assembly further comprises an inertiaelement coupled to and installed in the press fit element. The inertiaelement has a plurality of exterior helical splines corresponding to aplurality of interior helical splines of the press fit element.

The at least one translational inerter assembly further comprises atorsion bar having a torsion bar first end coupled to and installed inthe inertia element, and having a torsion bar second end fixedlyattached to the support structure of the aircraft. The torsion barfurther has a plurality of exterior linear splines corresponding to aplurality of interior linear splines of the inertia element. Rotation ofthe flight control surface causes translational movement of the inertiaelement, via the press fit element, along a hinge axis of the flightcontrol surface and along the torsion bar, resulting in the at least onetranslational inerter assembly damping movement of the flight controlsurface.

Also disclosed is a method for damping movement of a flight controlsurface of an aircraft. The method comprises the step of installing atleast one translational inerter assembly at a first end of the flightcontrol surface. The at least one translational inerter assemblycomprises a press fit element fixedly disposed within the first end ofthe flight control surface and rotatably movable with the flight controlsurface. The at least one translational inerter assembly furthercomprises an inertia element coupled to and installed in the press fitelement. The inertia element has a plurality of exterior helical splinescorresponding to a plurality of interior helical splines of the pressfit element. The at least one translational inerter assembly furthercomprises a torsion bar having a torsion bar first end coupled to andinstalled in the inertia element, and having a torsion bar second endfixedly attached to a support structure of the aircraft. The torsion barfurther has a plurality of exterior linear splines corresponding to aplurality of interior linear splines of the inertia element.

The method further comprises the step of rotating the flight controlsurface using one or more actuators. The method further comprises thestep of using the at least one translational inerter assembly to axiallyaccelerate in a translational movement along a hinge axis of the flightcontrol surface, the inertia element relative to the torsion bar,simultaneous with, and in proportion to, rotation of the flight controlsurface. The method further comprises the step of damping movement ofthe flight control surface, using the at least one translational inerterassembly.

Also disclosed is a translational inerter assembly for damping movementof a flight control surface of an aircraft with a support structure. Thetranslational inerter assembly comprises a press fit element rotatablydisposed within the flight control surface. The translational inerterassembly further comprises an inertia element disposed in the press fitelement. The translational inerter assembly further comprises a torsionbar coupled to the inertia element and to the support structure of theaircraft, such that when the flight control surface rotates, the inertiaelement translates, and movement of the flight control surface isdampened.

Also disclosed is an aircraft comprising a flight control surfacepivotably coupled to a support structure, one or more actuatorsactuating the flight control surface, and at least one translationalinerter assembly for damping movement of the flight control surface. Theat least one translational inerter assembly comprises a press fitelement rotatably disposed within the flight control surface. The atleast one translational inerter assembly further comprises an inertiaelement disposed in the press fit element. The at least onetranslational inerter assembly further comprises a torsion bar coupledto the inertia element and to the support structure, such that when theflight control surface rotates, the inertia element translates, andmovement of the flight control surface is dampened.

Also disclosed is a method for damping movement of a flight controlsurface of an aircraft having a support structure. The method comprisesthe step of attaching at least one translational inerter assembly to theflight control surface. The at least one translational inerter assemblycomprises a press fit element rotatably disposed within the flightcontrol surface, an inertia element disposed in the press fit element,and a torsion bar coupled to the inertia element and to the supportstructure of the aircraft. The method further comprises rotating theflight control surface using one or more actuators, to cause the inertiaelement, via the press fit element, to translate along a hinge axis ofthe flight control surface, and to translate along the torsion bar. Themethod further comprises damping movement of the flight control surface,using the at least one translational inerter assembly. The method mayfurther comprise, after attaching the at least one translational inerterassembly and prior to rotating the flight control surface, attachinganother translational inerter assembly to the flight control surface,wherein the two translational inerter assemblies comprise a firsttranslational inerter assembly, with a first press fit element, a firstinertia element, and a first torsion bar, attached to a first end of theflight control surface, and comprise a second translational inerterassembly, with a second press fit element, a second inertia element, anda second torsion bar, attached to a second end of the flight controlsurface.

The above-noted needs associated with actuators are specificallyaddressed and alleviated by the present disclosure which provides anapparatus including an inerter for damping an actuator. The inerterincludes a first terminal and a second terminal movable relative to oneanother along an inerter axis and configured to be mutually exclusivelycoupled to a support structure and a movable device actuated by anactuator. In one example, the inerter further includes a rod coupled toand movable with the first terminal. The inerter also includes athreaded shaft coupled to and movable with the second terminal. Theinerter additionally includes a flywheel having a flywheel annuluscoupled to the rod. The flywheel is configured to rotate in proportionto axial acceleration of the rod relative to the threaded shaft incorrespondence with actuation of the movable device by the actuator.

Also disclosed is an aircraft having a flight control surface pivotablycoupled to a support structure of the aircraft. The aircraft furtherincludes a hydraulic actuator configured to actuate the flight controlsurface. In addition, the aircraft includes an inerter having a firstterminal and a second terminal mutually exclusively coupled to thesupport structure and the flight control surface. The inerteradditionally includes a rod movable with the first terminal, and athreaded shaft movable with the second terminal. The inerter alsoincludes a flywheel coupled to the rod and the threaded shaft. Theflywheel is configured to rotate in proportion to axial acceleration ofthe rod relative to the threaded shaft in correspondence with actuationof the flight control surface by the actuator.

In addition, disclosed is a method of damping an actuator. The methodincludes actuating, using an actuator, a movable device. In addition,the method includes axially accelerating, using an inerter coupled tothe movable device, a first terminal relative to a second terminal ofthe inerter simultaneous with and in proportion to actuation of themovable device. Furthermore, the method includes rotationallyaccelerating a flywheel of the inerter in proportion to and simultaneouswith the axial acceleration of the first terminal relative to the secondterminal. Additionally, the method includes reducing actuator loadoscillatory amplitude of the movable device and actuator in response torotationally accelerating the flywheel.

The features, functions and advantages that have been discussed can beachieved independently in various examples of the present disclosure ormay be combined in yet other examples, further details of which can beseen with reference to the following description and drawings below.

BRIEF DESCRIPTION OF THE DRAWINGS

These and other features of the present disclosure will become moreapparent upon reference to the drawings wherein like numbers refer tolike parts throughout and wherein:

FIG. 1 is a block diagram of a flight control system of an aircraftincluding a hydraulic actuator for actuating a flight control surfaceand further including an inerter for damping the hydraulic actuator;

FIG. 2 is a block diagram of an example of an inerter integrated into ahydraulic actuator;

FIG. 3 is a perspective view of an aircraft;

FIG. 4 is a top view of a portion of a wing illustrating an actuator andan inerter operatively coupled to an aileron;

FIG. 5 is a sectional view of a wing taken along line 5 of FIG. 4 andillustrating an example of a linear hydraulic actuator mechanicallycoupled between a wing spar and one end of an aileron;

FIG. 6 is a sectional view of the wing taken along line 6 of FIG. 4 andillustrating an example of an inerter coupled to the aileron on an endopposite the actuator;

FIG. 7 is a sectional view of an example of a linear hydraulic actuatorhaving a piston axially slidable within an actuator housing;

FIG. 8 is a sectional view of an example of an inerter having a rodcoupled to a first terminal and a threaded shaft coupled to a secondterminal and including a flywheel threadably engaged to the threadedshaft and configured to rotate in proportion to axial acceleration ofthe rod and first terminal relative to the threaded shaft and secondterminal;

FIG. 9 is a magnified sectional view of the flywheel taken along line 9of FIG. 8 and illustrating a bearing rotatably coupling the flywheelannulus to the inerter rod and further illustrating the threadableengagement of the flywheel to the threaded shaft;

FIG. 10 is a sectional view of an example of an inerter integrated intoan unbalanced hydraulic actuator and illustrating the inerter flywheelrotatably coupled to a piston of the hydraulic actuator;

FIG. 11 is a sectional view of an example of an inerter having flywheelprotrusions for generating viscous damping within hydraulic fluid duringrotation of the flywheel;

FIG. 12 is a perspective view of an example of an inerter taken alongline 12 of FIG. 11 and illustrating a plurality of radially extendingflywheel blades circumferentially spaced around the flywheel perimeter;

FIG. 13 is a sectional view of an example of an inerter integrated intoa partially-balanced hydraulic actuator having an interior pistonaxially slidable within the piston rod;

FIG. 14 is a sectional view of an example of an inerter integrated intoa balanced hydraulic actuator having opposing piston sides withsubstantially equivalent cross-sectional areas;

FIG. 15 is a sectional view of an example of an inerter integrated intoa hydraulic actuator and wherein the flywheel is rotatably housed withinthe piston of the hydraulic actuator and including an electric flywheelmotor and a brake for actively controlling rotation of the flywheel;

FIG. 16 is a magnified sectional view of the flywheel and piston takenalong line 16 of FIG. 15 and illustrating the electric flywheel motorhaving permanent magnets mounted to the flywheel perimeter and windingsmounted to the piston inner wall;

FIG. 17 is a sectional view of an example of an inerter integrated intoa hydraulic actuator and wherein the flywheel and threaded shaft arerotatably coupled to the actuator end wall and the piston fixedlycoupled to the rod;

FIG. 18 is a magnified sectional view of the flywheel and piston takenalong line 18 of FIG. 17 and illustrating the flywheel annulus rotatablycoupled to the actuator end wall and the piston threadably engaged tothe threaded shaft in a manner such that linear translation of the rodrelative to the threaded shaft causes rotation of the flywheel andthreaded shaft;

FIG. 19 is a sectional view of an example of a flywheel rotatablycoupled to the actuator end wall and having an electric flywheel motorincluding permanent magnets mounted to the flywheel perimeter andwindings mounted to the housing side wall of the actuator;

FIG. 20 is a sectional view of a further example of a flywheel having anelectric flywheel motor and further including a brake configured toprovide dynamic braking of the flywheel;

FIG. 21 is a sectional view of an example of an inerter integrated intoa linear electro-mechanical actuator and illustrating the flywheelrotatably coupled to an actuator motor and threadably engaged to athreaded shaft;

FIG. 22 is a sectional view of an example of an inerter integrated intoa hydraulic actuator and illustrating the notations x, x₀, x₁, and x₂respectively denoting reference points for translation of the rod end,the cap end, the piston, and the flywheel wherein the notations are usedin the derivation of a transfer function characterizing the response ofan actuator having an integrated inerter;

FIG. 23 is a graph plotting frequency vs. magnitude (e.g., amplitude)for an actuator operating under a working pressure of 3000 psi, 5000psi, and 8000 psi, and illustrating a reduction in amplitude for theactuator damped by an inerter relative to the amplitude of the actuatorundamped by an inerter;

FIG. 24 is a flowchart having one or more operations that may beincluded in method of damping an actuator using an inerter;

FIG. 25 is a perspective view of an aircraft;

FIG. 26 is a top view of a wing section of a wing, taken along line26-26 of FIG. 25, illustrating an actuator and two translational inerterassemblies operatively coupled to a flight control surface in the formof an aileron;

FIG. 27 is a sectional view of the wing section, taken along line 27-27of FIG. 26, and illustrating an example of a first translational inerterassembly installed in a first end opening of a first end of the flightcontrol surface;

FIG. 28 is a sectional view of the wing section, taken along line 28-28of FIG. 26, and illustrating an example of a second translationalinerter assembly installed in a second end opening of a second end ofthe flight control surface;

FIG. 29A is a schematic perspective view of a flight control surfacewith a translational inerter assembly installed in a first end of theflight control surface;

FIG. 29B is a schematic perspective view of a flight control surfacewith two translational inerter assemblies installed the flight controlsurface with one installed in a first end and another installed in asecond end of the flight control surface;

FIG. 30A is an exploded perspective view of a flight control surfacewith a translational inerter assembly in a disassembled position;

FIG. 30B is a perspective view of the flight control surface of FIG. 30Awith the translational inerter assembly in an assembled position;

FIG. 31 is a block diagram of a flight control system of an aircraftincluding an actuator for actuating a flight control surface, andfurther including a translational inerter assembly for damping movementof the flight control surface; and

FIG. 32 is a flowchart having one or more operations that may beincluded in a method for damping movement of a flight control surface ofan aircraft.

DETAILED DESCRIPTION

Referring now to the drawings wherein the showings are for purposes ofillustrating various examples of the present disclosure, shown in FIG. 1is a block diagram of a hydraulic actuator 204 coupled between a supportstructure 116 and a movable device 124 and configured to move or actuatethe movable device 124. The block diagram advantageously includes arotational inerter 300 for damping the actuator 202. The inerter 300 isshown coupled between the support structure 116 and the movable device124 and is configured to improve the dynamic response of the movabledevice 124 during actuation by the actuator 202, as described in greaterdetail below. In the example shown in FIGS. 1 and 4-9, the inerter 300is provided as a separate component from the actuator 202. However, inother examples (e.g., FIGS. 2 and 10-21) described below, the inerter300 is integrated into the actuator 202.

The actuator 202 includes a piston 216 coupled to a piston rod 224. Thepiston 216 is slidable within an actuator housing 228 (e.g., acylinder). The actuator 202 further includes a rod end 214 and a cap end212 axially movable relative to one another in response to pressurizedhydraulic fluid acting in an unbalanced manner on one or both sides ofthe piston 216 inside the actuator housing 228. In the example shown,the rod end 214 is coupled to the movable device 124 and the cap end 212is coupled to the support structure 116. However, the actuator 202 maybe mounted such that the rod end 214 is coupled to the support structure116 and the cap end 212 is coupled to the movable device 124.

Referring still to FIG. 1, the inerter 300 includes a first terminal 302and a second terminal 304 axially movable or translatable relative toone another along an inerter axis 306 (FIG. 8) in correspondence withactuation of the movable device 124 by the actuator 202. In the exampleshown, the first terminal 302 is coupled to the movable device 124 andthe second terminal 304 is coupled to the support structure 116.However, the inerter 300 may be mounted such that the first terminal 302is coupled to the support structure 116 and the second terminal 304 iscoupled to the movable device 124. In an example not shown, the supportstructure to which the inerter 300 is coupled may be a different supportstructure than the support structure 116 to which the actuator 202 iscoupled.

The inerter 300 includes an inerter rod 308 coupled to and axiallymovable (e.g., translatable) with the first terminal 302. The inerterrod 308 may be aligned with or parallel to the inerter axis 306. Theinerter rod 308 may be hollow to define a rod bore 310. The threadedshaft 322 is coupled to and axially movable (e.g., translatable) withthe second terminal 304. The threaded shaft 322 may be aligned with orparallel to the inerter axis 306. The threaded shaft 322 has a free end324 that may be receivable within the rod bore 310. The threaded shaft322 may be hollow or may include a shaft bore 323 open on the free end324 of the threaded shaft 322. The threaded shaft 322 may include radialpassages 325 extending radially from the shaft bore 323 to the exteriorside of the threaded shaft 322 to allow fluid flow between the exteriorside of the threaded shaft 322 and the shaft bore 323. The shaft bore323 may allow fluid (e.g., hydraulic fluid—not shown) to flow from thefluid cavity at a second terminal 304 (for non-integratedinerters—FIG. 1) or cap end 212 (for integrated inerters—FIG. 2),through the shaft bore 323, and into the fluid cavity at the flee end324 (FIG. 8) of the threaded shaft 322 to allow the fluid to lubricatemoving parts of the bearing 328 and/or at the flywheel annulus 318. Thesize (e.g., diameter) of the shaft bore 323 and the size (e.g.,diameter) and quantity of the radial passages 325 may be configured toapportion fluid flow to the bearing 328 and the flywheel annulus 318.

As shown in FIG. 1, the inerter 300 includes a flywheel 314 (e.g., aspinning mass). In some examples (e.g., FIGS. 6 and 8-16), the flywheel314 is threadably coupled to the threaded shaft 322 which convertslinear motion of the threaded shaft 322 into rotational motion of theflywheel 314. The flywheel 314 is configured to rotate in proportion toaxial movement of the inerter rod 308 relative to the threaded shaft 322in correspondence with actuation of the movable device 124 by theactuator 202. In this regard, the flywheel 314 is configured torotationally accelerate and decelerate in proportion to axialacceleration and deceleration of the inerter rod 308 (e.g., coupled tothe first terminal 302) relative to the threaded shaft 322 (e.g.,coupled to the second terminal 304).

Advantageously, the flywheel 314 is coupled to the inerter rod 308 at aflywheel annulus 318 and is threadably engaged to the threaded shaft322, as shown in FIGS. 1, 8-9, and 14 and described in greater detailbelow. However, in other examples, the flywheel annulus 318 may becoupled to the piston 216 as shown in FIGS. 10-13 and 15-16 anddescribed below. In still further examples, the flywheel annulus 318 maybe coupled to the actuator housing 228 as shown in FIGS. 17-20 anddescribed below.

Regardless of the component to which the flywheel 314 is coupled, theflywheel 314 may include at least one bearing 328 (e.g., a thrustbearing 328) at the flywheel annulus 318 to rotatably couple theflywheel 314 to the inerter rod 308 (FIGS. 1, 8-9, and 14), the piston216 (FIGS. 10-13 and 15-16), or the actuator housing 228 (FIGS. 17-20).The bearing 328 allows the flywheel 314 to axially translate with theinerter rod 308 as the flywheel 314 rotates on the threads of thethreaded shaft 322 in response to axial movement of the inerter rod 308relative to the threaded shaft 322. Advantageously, by coupling theflywheel 314 to the component (i.e., the inerter rod 308, the piston216, or the actuator housing 228) at the flywheel annulus 318 instead ofat the flywheel perimeter 316, the flywheel 314 exhibits limited flexurein the axial direction during high-frequency, oscillatory, axialacceleration of the first terminal 302 relative to the second terminal304. Such axial flexure of the flywheel mass would otherwise reduceflywheel rotational motion during high-frequency, oscillatory, axialacceleration.

Referring still to the example of FIG. 1, the support structure 116 isshown configured as a wing spar 118 of a wing 114 of an aircraft 100.The movable device 124 is shown as a flight control surface 122 of aflight control system 120 of the aircraft 100. The flight controlsurface 122 may be hingedly coupled to the rigid support structure 116such as a wing spar 118 or other structure. The flight control surface122 may be pivotably about a hinge axis 126. The flight control surface122 may comprise any one of a variety of different configurationsincluding, but not limited to, a spoiler, an aileron, an elevator 112,an elevon, a flaperon, a rudder 108, a high-lift device such as aleading edge slat, a trailing edge flap, or any other type of movabledevice 124.

The actuator 202 provides positive force to move the flight controlsurface 122 to a commanded position in response to a command input fromthe flight crew or an autopilot. The inerter 300 provides for controland damping of displacements of the flight control surface 122. One ormore inerters 300 may be included in a flight control system 120. In oneexample, the one or more inerters 300 may be configured to suppress orprevent control surface flutter as may be aerodynamically-induced at aresonant frequency of the flight control surface 122. For example, thepresently-disclosed inerter 300 may be configured to reduce actuatorload oscillatory amplitude at resonance (e.g., at a resonant frequency)of up to approximately 20 Hz (e.g., ±5 Hz) which may correspond to theflutter frequency of a flight control surface 122 of an aircraft 100.Additionally or alternatively, the inerter 300 may provide additionalfunctionality for improving the dynamic response of a movable device124, such as increasing the actuation rate of the movable device 124and/or preventing position overshoot of a commanded position of themovable device 124, as described in greater detail below.

In one example, the inerter 300 may be configured such that rotation ofthe flywheel 314 reduces actuator load oscillatory amplitude atresonance of the coupled actuator 202 and movable device 124 by at leastapproximately 10 percent relative to the actuator load oscillatoryamplitude that would otherwise occur using the same actuator 202 withoutan inerter 300. Advantageously, the presently-disclosed inerter 300permits the operating bandwidth of the actuator 202 to encompass ormatch the resonant frequency of the coupled movable device 124 andactuator 202 without the potential for oscillatory response, without thepotential for exceeding the strength capability of the mounting system(not shown) of the flight control surface 122 and actuator 202, and/orwithout the potential for flight control surface 122 deflections thatcould aerodynamically destabilize the aircraft 100.

The presently-disclosed examples of the inerter 300 allow for areduction in the overall size and weight of an actuator 202 systemwithout the potential for oscillatory response. More specifically, theinerter 300 allows for a reduction in the inertia or inertial load onthe actuator 202 which, in turn, allows for a reduction in pistoncross-sectional area of the actuator 202 and a decrease in the size andweight of other hydraulic system components including reservoirs, tubingdiameter, accumulators, pumps, and other components. In this regard, theinerter 300 increases the power density for a hydraulic actuator systemin any application where dynamic response is limited by pistoncross-sectional area or load inertia. The presently-disclosed inerter300 examples may be implemented with hydraulic actuators 204 configuredto be operated at a working pressure of at least 5000 psi. For example,the inerter 300 examples may be implemented with hydraulic actuators 204operated at a working pressure of approximately 3000 psi and, in someexamples, the hydraulic actuators 204 may be operated at a workingpressure of approximately 8000 psi. A relatively high working pressureof a hydraulic actuator 204 may facilitate a reduction in total flow ofhydraulic fluid through the hydraulic system (e.g., flight controlsystem 120) which may enable a reduction in the volumetric requirementfor hydraulic fluid reservoirs and accumulators.

In the case of an aircraft 100, the reduced size of the actuators 202may reduce the amount by which such actuators 202 protrude outside ofthe outer mold line (not shown) of the aircraft 100 with a resultingdecrease in aerodynamic drag. Even further, the presently-disclosedinerter examples may allow for a reduction in the amount of off-takepower from the aircraft propulsion units (e.g., gas-turbine engines)which may provide the potential for using higher bypass ratio gasturbine engines such as in commercial aircraft applications. Thedecrease in the size of the hydraulic system, the reduction inaerodynamic drag, and/or the reduction in off-take power may translateto an increase in aircraft performance including, but not limited to,increased fuel efficiency, range, and/or payload capacity.

Although the presently-disclosed inerter examples are described in thecontext of a linear hydraulic actuator 204, the inerter 300 may beimplemented in other types of actuators 202 including, but not limitedto, a rotary hydraulic actuator, an electro-hydraulic actuator (e.g.,rotary or linear), a mechanical actuator, an electro-mechanicalactuator, and other types of actuators. In one example (see FIG. 21),the electro-mechanical actuator 242 may be a linear electro-mechanicalactuator having a threaded shaft 322 coupled to a movable device 124. Asdescribed in greater detail below with reference to FIG. 21, the linearelectro-mechanical actuator 242 may include an electric actuator motor244 for causing axial motion of a threaded shaft 322. A flywheel 314 maybe threadably engaged to the threaded shaft 322 and may be configured torotationally accelerate and decelerate in proportion to axialacceleration and deceleration of the threaded shaft 322 during actuationof the movable device 124 by the linear electro-mechanical actuator 242.

It should also be noted that although the presently-disclosed inerterexamples are described in the context of an aircraft flight controlsystem 120, any one of the inerters 300 may be implemented in any typeof open-loop or closed-loop control system for use in any one of avariety of different applications in any industry, without limitation.In this regard, the presently-disclosed inerters 300 may be implementedin any vehicular application or non-vehicular application. For example,an inerter 300 may be implemented in any marine, ground, air, and/orspace application, and in any vehicular or non-vehicular system,subsystem, assembly, subassembly, structure, building, machine, andapplication that uses an actuator to actuate a movable device.

In some examples, an inerter 300 may be implemented for damping movementof a movable device configured to control the direction of travel of avehicle. For example, an inerter may be implemented for damping movementof aerodynamic control surfaces of an air vehicle, hydrodynamic controlsurfaces of a marine vessel, thrust directors including thrust-vectoringnozzles of an aircraft or a launch vehicle (e.g., a rocket), or anyother type of mechanical device that influences the direction of travelof a vehicle and which may be susceptible to external vibratory forces.In a specific example of a wheeled vehicle configured to move over land,any one of the presently-disclosed inerter examples may be implementedin a steering system to control or avoid wheel shimmy, such as may occurin a steerable wheel of an aircraft landing gear such as a nose landinggear.

FIG. 2 is a block diagram of an example of an inerter 300 integratedinto a hydraulic actuator 204 coupled between a support structure 116and a flight control surface 122 of a flight control system 120 of anaircraft 100. In the example shown, the actuator 202 is a linearhydraulic actuator 204 having a piston 216 coupled to a rod (e.g.,piston rod 224) and axially slidable within a housing (not shown). Inthe example shown, the flywheel 314 of the inerter 300 is rotatablycoupled to the piston 216 at the flywheel annulus 318. The flywheel 314is threadably coupled to the threaded shaft 322 and configured torotationally accelerate in proportion to axial acceleration of thepiston 216 and rod relative to the threaded shaft 322. However, asmentioned above, the flywheel 314 may be rotatably coupled to the piston216 (e.g., FIGS. 10-16) or the flywheel 314 may be rotatably coupled tothe cap end 212 (e.g., FIGS. 17-20) or rod end 214 of the actuatorhousing 228.

As mentioned above, the threaded shaft 322 may include a shaft bore 323open on the free end 324 and having radial passages 325 to allow fluid(e.g., hydraulic fluid) to flow from the cap end chamber 236 at the capend 212), through the shaft bore 323, and out of the free end 324 of thethreaded shaft 322 to allow the fluid to lubricate moving parts of thebearing 328 and/or the flywheel annulus 318. The shaft bore 323 andradial passages 325 may be included in any one of the inerter 300examples disclosed herein.

In the present disclosure, for examples wherein the inerter 300 isintegrated into the actuator 202, the rod end 214 or cap end 212 of theactuator 202 functions as the first terminal 302 of the inerter 300, andthe remaining rod end 214 or cap end 212 of the actuator 202 functionsas the second terminal 304 of the inerter 300. In this regard, the terms“first terminal” and “second terminal” are non-respectively usedinterchangeably with the terms “rod end” and “cap end.” In addition, forexamples where the inerter 300 is integrated into the actuator 202, theterm “rod” is used interchangeably with the terms “piston rod” and“inerter rod.” Similarly, for examples where the inerter 300 isintegrated into the actuator 202, the term “housing” is usedinterchangeably with the terms “actuator housing” and “inerter housing.”

FIG. 3 is a perspective view of an aircraft 100 having one or moreinerters 300 for control and/or damping of one or more actuators 202.The aircraft 100 may include a fuselage 102 and a pair of wings 114extending outwardly from the fuselage 102. The aircraft 100 may includea pair of propulsion units (e.g., gas turbine engines). As mentionedabove, each wing 114 may include one or more movable devices 124configured as flight control surfaces 122 which may be actuated by anactuator 202 damped and/or assisted by an inerter 300. Such flightcontrol surfaces 122 on the wings 114 may include, but are not limitedto, spoilers, ailerons, and one or more high-lift devices such as aleading edge slats and/or trailing edge flaps. At the aft end of thefuselage 102, the empennage 104 may include one or more horizontal tails110 and a vertical tail 106, any one or more of which may include flightcontrol surfaces 122 such as an elevator 112, a rudder 108, or othertypes of movable devices 124 that may be actuated by an actuator 202damped and/or assisted by an inerter 300.

FIG. 4 is a top view of a portion of the wing 114 of FIG. 3 illustratingan aileron actuated by a hydraulic actuator 204 located on one end ofthe aileron and having an inerter 300 located on an opposite and theaileron 130. The aileron 130 may be hingedly coupled to a fixed supportstructure 116 of the wing 114 such as a wing spar 118. In FIG. 4, thehydraulic actuator 204 and the inerter 300 are provided as separatecomponents and may each be coupled between the support structure 116(e.g., the wing spar 118) and the aileron 130.

FIG. 5 is a sectional view of the wing 114 of FIG. 4 showing an exampleof a linear hydraulic actuator 204 mechanically coupled between the wingspar 118 and one end of the aileron 130. In the example shown, the rodend 214 of the hydraulic actuator 204 is coupled to a bellcrank 128. Thebellcrank 128 is hingedly coupled to the aileron in a manner such thatlinear actuation of the hydraulic actuator 204 causes pivoting of theaileron about the hinge axis 126. The cap end 212 of the hydraulicactuator 204 is coupled to the wing spar 118.

FIG. 6 is a sectional view of the wing 114 of FIG. 4 and showing anexample of an inerter 300 coupled between the wing spar 118 and theaileron 130. As mentioned above, the inerter 300 is located on an end ofthe aileron opposite the hydraulic actuator 204. The first terminal 302of the inerter 300 is coupled to a bellcrank 128. The second terminal304 of the inerter 300 is coupled to the wing spar 118. Due to thehydraulic actuator 204 and the inerter 300 being coupled to the samemovable device 124 (i.e., the aileron 130), relative axial accelerationof the cap end 212 and rod end 214 of the actuator 202 causesproportional axial acceleration of the first terminal 302 and secondterminal 304 of the inerter 300 resulting in rotational acceleration ofthe flywheel 314.

FIG. 7 is a partially cutaway sectional view of an example of adouble-acting hydraulic actuator 204 having a cap end 212 and a rod end214 axially movable relative to one another during actuation of themovable device 124. As mentioned above, the rod end 214 and the cap end212 may be mutually exclusively coupled to the support structure 116 andthe movable device 124. For example, the rod end 214 may be coupled tothe support structure 116 and the cap end 212 may be coupled to themovable device 124, or the rod end 214 may be coupled to the movabledevice 124 and the cap end 212 may be coupled to the support structure116.

In FIG. 7, the piston 216 is coupled to a free end 324 of the piston rod224 and is axially slidable within the actuator housing 228. The piston216 divides the actuator housing 228 into a cap end chamber 236 and arod end chamber 238. The actuator housing 228 of the double-actinghydraulic actuator 204 includes a pair of fluid ports 234 through whichpressurized hydraulic fluid enters and leaves the cap end chamber 236and the rod end chamber 238 chambers for moving the piston 216 withinthe actuator housing 228. In any of the presently-disclosed examples,the hydraulic actuator 204 may also be configured as a single-actingactuator (not shown) wherein the actuator housing 228 contains a singlefluid port 234 for receiving pressurized hydraulic fluid in the actuatorhousing 228 as a means to move the piston 216 along one direction withinthe actuator housing 228, and optionally include a biasing member (e.g.,a spring—not shown) for moving the piston 216 in an opposite direction.

FIG. 8 is a partially cutaway sectional view of an example of an inerter300 having an inerter housing 330 containing the flywheel 314 and havingan inerter side wall 334 and opposing inerter end walls 332. One inerterend wall 332 may include a housing bore through which the inerter rod308 extends and terminates at the first terminal 302. The inerter 300includes a threaded shaft 322 coupled to the inerter end wall 332located at the second terminal 304. In the example of FIG. 8, theflywheel 314 is coupled to an end of the inerter rod 308 and threadablyengaged to the threaded shaft 322. The flywheel 314 rotates inproportion to axial acceleration of the inerter rod 308 and firstterminal 302 relative to the threaded shaft 322 and second terminal 304.

FIG. 9 is a magnified sectional view of FIG. 8 showing the flywheel 314coupled to the inerter rod 308 at the flywheel annulus 318. The flywheelannulus 318 is also threadably engaged to the threaded shaft 322. In theexample shown, the threaded shaft 322 is configured as a ball screw 326having helical grooves for receiving ball bearings which couplesimilarly-configured helical grooves in the flywheel annulus 318 to theball screw 326 with minimal friction. Although not shown, the flywheelannulus 318 may include a ball nut for circulating the ball bearingscoupling the flywheel 314 to the ball screw 326. In another example notshown, the threaded shaft 322 may comprise a lead screw having threadsto which the flywheel annulus 318 are directly engaged. As may beappreciated, the flywheel 314 may be configured for engagement to anyone of a variety of different types of configurations of threadedshafts, and is not limited to the ball screw 326 example illustrated inFIG. 9.

Also shown in FIG. 9 is an example of a bearing 328 for coupling theflywheel annulus 318 to the inerter rod 308 such that the inerter rod308 and flywheel 314 may translate in unison as the flywheel 314 rotatesdue to threadable engagement with the threaded shaft 322. Although thebearing 328 is shown as a ball bearing, the bearing 328 may be providedin any one a variety of different configurations capable of axiallycoupling the flywheel 314 to the inerter rod 308 with a minimal amountof axial free play. For example, the bearing 328 may be configured as aroller bearing (not shown). In still further examples, the flywheel 314may be coupled to the inerter rod 308 without a bearing while stillallowing the flywheel 314 to rotate during translation of the inerterrod 308 and flywheel 314 relative to the threaded shaft 322.

FIG. 10 is a sectional view of an example of an inerter 300 integratedinto a hydraulic actuator 204 having a housing containing a piston 216.The actuator 202 is a double-acting actuator including a pair of fluidports 234 for receiving pressurized hydraulic fluid in a cap end chamber236 and a rod end chamber 238 located on opposite sides of the piston216. The actuator 202 is an unbalanced actuator 206 wherein one of thepiston sides 218 has a greater cross-sectional area than the oppositepiston side 218. The piston 216 may include a piston 216 seal (e.g., anO-ring seal—not shown) extending around the piston perimeter 220 forsealing the piston perimeter 220 to the actuator side wall 232.

As mentioned above, for examples where the inerter 300 is integratedinto an actuator 202, the rod end 214 or the cap end 212 of the actuator202 functions as the first terminal 302 of the inerter 300, and theremaining rod end 214 or the cap end 212 of the actuator 202 functionsas the second terminal 304 of the inerter 300. In the example shown, theflywheel 314 is mounted in the cap end chamber 236 and is rotatablycoupled to the piston 216 at the flywheel annulus 318. The flywheel 314is threadably engaged to the threaded shaft 322 which passes through thepiston 216 and extends into the rod bore 310. The flywheel 314 isconfigured to rotationally accelerate in proportion to axialacceleration of the piston 216 and piston rod 224 relative to thethreaded shaft 322.

FIG. 11 shows an example of an inerter 300 having flywheel protrusions320 for generating viscous damping during rotation of the flywheel 314when the flywheel 314 is immersed in hydraulic fluid. The flywheelprotrusions 320 generate or increase the viscous damping capability ofthe inerter 300 during rotation of the flywheel 314, and therebyincrease the damping capability of the inerter 300.

FIG. 12 is a perspective view of an example of an inerter 300 having aplurality of radially extending flywheel blades circumferentially spacedaround the flywheel perimeter 316. During rotation of the flywheel 314,the flywheel blades may generate viscous damping capability and add tothe inerting capability of the inerter 300. Although FIG. 12 illustratesthe flywheel protrusions 320 as radially-extending flywheel blades, theflywheel 314 may be provided with flywheel protrusions 320 extendingfrom any portion of the flywheel 314 including one or both of theopposing sides of the flywheel 314. In addition, the flywheelprotrusions 320 may be provided in any geometric size, shape orconfiguration, without limitation, and are not limited to flywheelblades.

FIG. 13 is a sectional view of an example of an inerter 300 integratedinto a hydraulic actuator 204 configured as a partially-balancedactuator 208. The partially-balanced actuator 208 includes an interiorpiston 226 coupled to a free end 324 of the threaded shaft 322. Theinterior piston 226 may be axially slidable within the rod bore 310 andmay be rotatably coupled to the end of the threaded shaft 322 such thatthe interior piston 226 is non-rotatable relative to the rod bore 310during axial movement of the piston rod 224 relative to the threadedshaft 322. Although not shown, the perimeter of the interior piston 226may be sealed (e.g., via an O-ring) to the rod wall 312 of the rod bore310. The inclusion of the interior piston 226 may reduce the totalvolume of hydraulic fluid required to fill the cap end chamber 236during extension of the piston rod 224 relative to the increased volumeof hydraulic fluid required to fill the cap end chamber 236 for examples(e.g., FIG. 8) lacking an interior piston 226.

FIG. 14 is a partially cutaway sectional view of an example of aninerter 300 integrated into a hydraulic actuator 204 configured as abalanced actuator 210 having opposing piston sides 218 withsubstantially equivalent cross-sectional areas. The housing may includea separator wall 240 separating the portion of the housing containingthe flywheel 314 from the portion of the housing containing the piston216. A cap end chamber 236 is located on one of the piston sides 218 andthe rod end chamber 238 is located on the opposite piston side 218. Thepiston 216 may be mounted on the piston rod 224. In FIG. 14, one end ofthe piston rod 224 extends through the actuator end wall 230 andterminates at the rod end 214 (e.g., the first terminal 302). Anopposite end of the piston rod 224 extends through the separator wall240. The flywheel 314 is rotatably coupled to the piston rod 224 in amanner as described above.

FIG. 15 is a partially cutaway sectional view of an example of aninerter 300 having an electric flywheel motor 350 integrated into ahydraulic actuator 204. The flywheel motor 350 may facilitate activecontrol of flywheel 314 rotation using electromotive force from theintegrated flywheel motor 350. Active control may include using theflywheel motor 350 to apply a torque to the flywheel 314 to resist oraid the torque that is generated by the flywheel 314 due to axialacceleration of the first terminal 302 relative to the second terminal304. The flywheel motor 350 may be configured to provide active dampingand/or active braking of the actuator 202 and the load inertia.

FIG. 16 is a magnified sectional view of FIG. 15 showing the flywheel314 rotatably coupled to and contained within a generally hollow piston216 which is actually slidable within the actuator housing 228. Alsoshown in the flywheel motor 350 incorporated into the flywheel 314 andthe piston 216 and configured to actively control rotation of theflywheel 314 in correspondence with relative axial movement of the rodand threaded shaft 322. The flywheel motor 350 may be operated in amanner to accelerate and/or decelerate the flywheel 314 by applying atorque to the flywheel 314 either in correspondence with (e.g., the samedirection as) or in opposition to the direction of rotation of theflywheel 314. In this manner, the flywheel motor 350 may apply a torqueto the flywheel 314 to resist or aid the flywheel torque generated dueto axial acceleration of the first terminal 302 relative to the secondterminal 304.

In the example of FIG. 16, the flywheel motor 350 is a permanent magnetdirect-current (DC) motor having one or more permanent magnets 354mounted to the flywheel 314. For example, a plurality of permanentmagnets 354 may be circumferentially spaced around the flywheelperimeter 316. In addition, the flywheel motor 350 may include aplurality of windings 352 mounted to the piston 216. In one example, aplurality of windings 352 may be circumferentially spaced around thepiston inner wall 222 (e.g., FIGS. 15-16). In another example, aplurality of windings 352 may be circumferentially spaced around theside wall 232 of the housing (e.g., FIGS. 19-20) as described below. Inother examples, the flywheel motor 350 may be a brushless DC motor orsome other motor configuration, and is not limited to a permanent magnetDC motor configuration as shown in FIGS. 15-16 and 19-20. In an examplenot shown, a linear position sensor may be included with the actuator202 to sense the linear position of the piston 216 and generate a signalrepresentative of the linear piston position for commutating theflywheel motor 350 in correspondence with the piston position.

As mentioned above, the flywheel motor 350 in FIGS. 15-16 may beconfigured to assist or aid in rotating the flywheel 314 for a commandeddirection of motion of the movable device 124. For example, the flywheelmotor 350 may provide a torque to accelerate the flywheel 314 at thestart of motion of the movable device 124 toward a commanded position.The torque applied to the flywheel 314 by the flywheel motor 350 may beapproximately equal in magnitude to the torque required to rotationallyaccelerate the flywheel 314 due to axial acceleration of the threadedshaft 322 relative to the rod. By using the flywheel motor 350 to removethe torque required to rotationally accelerate the flywheel 314, thepiston 216 may move more quickly to a commanded position than if theflywheel motor 350 did not accelerate the flywheel 314. In this manner,the flywheel motor 350 may allow faster responsiveness of a movabledevice 124 than a conventional actuator 202. The level of dampingprovided by an inerter 300 having active control of the flywheel 314 maybe greater than the damping that is feasible in a closed-loop controlsystem without active control due to the risk of control systeminstability. Although FIGS. 15-16 illustrate a flywheel motor 350incorporated into an inerter 300 integrated with an actuator 202, aflywheel motor 350 may be incorporated into an inerter 300 that is aseparate component from the actuator 202 (e.g., FIGS. 4-8).

In a further example of active control, the flywheel motor 350 may beoperated in a manner to provide a torque to decelerate the flywheel 314as the movable device 124 approaches a commanded position. In thisregard, the flywheel motor 350 may be operated as a brake to oppose theflywheel torque generated by the axial deceleration of the threadedshaft 322 relative to the piston rod 224. Actively controlling flywheel314 rotation in this manner may prevent or limit position overshoot ofthe movable device 124 and thereby increase the stability of the movabledevice 124. In such an arrangement, the actuator 202 and inerter 300 maybe configured with a failure mode that ensures that without active motorcontrol, the actuator 202 is capable of exhibiting a desired dampedresponse in a manner preventing underdamping of the movable device 124.An inerter 300 having a flywheel motor 350 for active control may beconnected to the movable device 124 without being part of the actuator202 such that in the event of a disconnect of the actuator 202 from themovable device 124 or in the event of a failure of the actuator 202 tohold the load of the movable device 124, the flywheel motor 350 may beoperated in a manner preventing underdamped movement of the movabledevice 124 for the given failure mode.

Referring still to FIG. 16, in another example of active control, theflywheel motor 350 may include a brake 360 configured to provide dynamicbraking of the flywheel 314. In this regard, the brake 360 may beoperated in a manner to decelerate the flywheel 314 or to increaseexisting deceleration of the flywheel 314. For examples that include aflywheel motor 350, the brake 360 may be operated in a manner toincrease existing deceleration of the flywheel 314 caused by rotationaldrag of the flywheel motor 350. In addition, the flywheel motor 350 maybe operated in a manner to oppose disturbances (e.g., undesirablemotion) of the actuator 202.

In the example of FIG. 16, the brake 360 may be configured as a discbrake having brake pads 364. The flywheel 314 may function as a brakerotor against which the brake pads 364 may be frictionally engagedduring braking. In other examples not shown, a separate brake rotor maybe provided which may be directly or indirectly coupled to the flywheel314. In the example shown, a hydraulic brake cylinder (not shown) may beincluded to actuate the brake pads 364 into frictional engagement withone or both of the opposing axial faces 362 (e.g., planar faces) of theflywheel 314 for decelerating the flywheel 314. Preferably, the brake360 may include at least two pairs of opposing brake pads 364 located ondiametrically opposing sides of the brake rotor. Each pair of brake pads364 may be held in position by a bracket 366. Although the brake 360 isdescribed and illustrated as a disc brake, the inerter 300 mayincorporate any one or more different types of brakes such as a drumbrake or any other type of brake capable of decelerating the flywheel314.

Referring to FIG. 17, shown is a partially cutaway sectional view ofanother example of an inerter 300 integrated into a hydraulic actuator204. The flywheel 314 is rotatably coupled or attached to the actuatorend wall 230 which may be coupled to the second terminal 304. The piston216 is fixedly coupled or attached to the piston rod 224 which extendsfrom the piston 216 through the actuator end wall 230 and is coupled tothe first terminal 302. In an alternative example not shown, theflywheel 314 may be rotatably coupled to the actuator end wall 230 whichis attached to the first terminal 302, and the piston rod 224 may becoupled to the second terminal 304.

FIG. 18 is a magnified sectional view of FIG. 17 illustrating theflywheel annulus 318 rotatably coupled by a bearing 328 to the actuatorend wall 230. The threaded shaft 322 is fixedly coupled to the flywheel314 and is rotatable in unison with the flywheel 314. As mentionedabove, the piston 216 is fixedly coupled to the piston rod 224 andthreadably engaged to the threaded shaft 322 in a manner such thatlinear translation of the piston rod 224 relative to the threaded shaft322 causes rotation of the flywheel 314 and threaded shaft 322 inunison. As indicated above, axial movement of the threaded shaft 322relative to the piston rod 224 may be in correspondence with actuationof the movable device 124 by the actuator 202.

FIG. 19 illustrates an example of a flywheel 314 rotatably coupled tothe actuator end wall 230 and incorporating a flywheel motor 350 foractive control of the rotation of the flywheel 314 in a manner asdescribed above. The flywheel motor 350 may include permanent magnets354 mounted to the flywheel perimeter 316. For example, as describedabove with regard to FIG. 16, a plurality of permanent magnets 354 maybe circumferentially spaced around the flywheel perimeter 316. FIG. 19also shows a plurality of windings 352 circumferentially spaced aroundthe actuator side wall 232 of the actuator housing 228.

FIG. 20 illustrates an example of a flywheel 314 including a brake 360configured to provide dynamic braking of the flywheel 314. In theexample shown, the brake 360 is configured as a disc brake having one ormore pairs of brake pads 364 for frictionally engaging opposing axialfaces 362 of the flywheel 314. The brake 360 in FIG. 20 may beconfigured and operated similar to the arrangement illustrated in FIG.16 and described above.

FIG. 21 illustrates an example of an inerter 300 integrated into alinear electro-mechanical actuator 242. The electro-mechanical actuator242 may extend between a support structure 116 (FIG. 2) and a movabledevice 124 (FIG. 2). The electro-mechanical actuator 242 may include anelectric actuator motor 244 supported by the actuator housing 228. Thefirst terminal 302 may be coupled to a movable device 124. Theelectro-mechanical actuator 242 may include a second terminal 304 whichmay be coupled to a support structure 116. Alternatively, the firstterminal 302 may be coupled to the support structure 116 and the secondterminal 304 may be coupled to the movable device 124.

The electro-mechanical actuator 242 may include a threaded shaft 322(e.g., an Acme-threaded shaft, a ball screw, etc.) extending through theactuator motor 244 and terminating at the first terminal 302. Theactuator motor 244 may be operably coupled to the threaded shaft 322 bya motor-shaft coupler 246 which may be threadably engaged to thethreaded shaft 322. Operation of the actuator motor 244 may cause axialmotion of the threaded shaft 322 for actuating the movable device 124.In this regard, the threaded shaft 322 may axially move in proportion(e.g., in magnitude and direction) to angular displacement of theactuator motor 244. A flywheel 314 may be threadably engaged to thethreaded shaft 322. In addition, the flywheel annulus 318 may berotatably coupled to the actuator motor 244 via a bearing 328 such thataxial acceleration of the threaded shaft 322 causes rotationalacceleration of the flywheel 314. The flywheel 314 may be configured torotationally accelerate and decelerate in proportion to axialacceleration and deceleration of the threaded shaft 322 (e.g., relativeto the actuator motor 244) during actuation of the movable device 124.

In this regard, rotation of the flywheel 314 during actuation of theelectro-mechanical actuator 242 of FIG. 21 may provide any one or moreof the advantages described herein for improving the dynamic response ofthe movable device 124 during actuation by the electro-mechanicalactuator 242. For example, the flywheel 314 may reduce actuator loadoscillatory amplitude at resonance of the coupled electro-mechanicalactuator 242/movable device 124. In addition, although not shown in FIG.21, a flywheel motor 350 (e.g., FIG. 16) and/or a dynamic brake 360(FIG. 16) may optionally be included with the flywheel 314 to allow foractive control of the rotation of the flywheel 314 using any one or moreof the flywheel control techniques described herein.

FIG. 22 is a sectional view of an example of an inerter 300 integratedinto a hydraulic actuator 204 as described above and illustrated in FIG.10. FIG. 22 includes the notations x, x₀, x₁, and x₂ respectivelydenoting reference points for translation of the rod end 214, the capend 212, the piston 216, and the flywheel 314. The notations x, x₀, x₁,and x₂ are parameters that are used in a below-described of a transferjunction

$\frac{X(s)}{F(s)}$(Equation 220) mathematically characterizing the response of theapparatus of FIG. 22. Table 1 includes a listing of the parameters usedin the derivation of the transfer function. Included with each listedparameter is an indication of the physical type of the parameter and abrief description of the parameter.

TABLE 1 Parameter Physical type Description F force actuator reactedforce (newton) F₁ force piston reacted force (newton) F₂ force flywheelreacted force (newton) F₃ force flywheel to piston compliance force(newton) T₂ torque flywheel acceleration torque (newton-meter) {dot over(x)} translation actuator rod end translation reference (meter) x₁, {dotover (x)}₁, {umlaut over (x)}₁ translation piston translation reference(meter) x₂, {dot over (x)}₂, {umlaut over (x)}₂ translation flywheeltranslation transformed from rotation (meter) x₀, {dot over (x)}₀,{umlaut over (x)}₀ translation actuator cap end translation reference(meter) θ, {dot over (θ)}, {umlaut over (θ)} rotation flywheel rotationreference (radian) J mass moment of flywheel inertia in rotation(kilogram-meter²) inertia B damping coefficient flywheel torqueresisting {dot over (θ)} (newton-meter/radian/s) Z stiffness flywheelrotational stiffness (radian/newton-meter) M mass actuator reactedinertia at rod end (kilogram) C damping coefficient actuator forceresisting {dot over (x)} (newton/meter/s) K stiffness actuator stiffness(meter/newton) r thread rate revolutions per translation (radian/meter)ω_(n) natural frequency 2^(nd) order model characteristic(radian/second) ζ damping factor 2^(nd) order model characteristic (nounit) c constant zero offset to a common reference

Equations 100 to 210 inclusive are the assumptions behind the derivationof the transfer function of Equation 220. Referring to the exampleapparatus of FIG. 22, the total reacted force F (e.g., at the rod end214) may be computed as the sum of the piston 216 reacted force F₁ andthe flywheel 314 reacted force F₂ as shown in Equation 100, wherein thesign of F₁ and F₂ are the same from a disturbance rejection sense:F=F ₁ +F ₂  (Equation 100)

The torque T₂ developed by the flywheel 314 may be determined usingEquation 110 as the sum of the product of the flywheel rotationalinertia J and flywheel rotational acceleration {umlaut over (θ)} and theproduct of a flywheel damping coefficient B and the flywheel rotationalvelocity {dot over (θ)}:T ₂ =J{umlaut over (θ)}+B{dot over (θ)}  (Equation 110)

The flywheel reacted force F₂ may be computed using equation 120 as theproduct of the flywheel torque T₂ and the thread rate r (e.g., threadpitch) of the threaded shaft 322. The thread rate may be described asthe linear distance of travel of the flywheel 314 per revolution:F ₂ =r(J{umlaut over (θ)}+B{dot over (θ)})  (Equation 120)

The rotation of the flywheel 314 may be characterized by the flywheelangular displacement or rotational angle θ, rotational velocity {dotover (θ)}, and rotational acceleration {umlaut over (θ)}, asrespectively represented by Equations 130, 140, and 150. The flywheelrotational angle θ is the product of the thread rate r and the lineardistance of flywheel translation x₂ as represented by Equation 130. Theparameter c is a constant representing a linear offset relative to acommon reference. The flywheel rotational velocity {dot over (θ)} is theproduct of the thread rate r and the linear velocity {dot over (x)}₂ ofthe flywheel 314 as represented by Equation 140. The flywheel rotationalacceleration {umlaut over (θ)} is the product of the thread rate r andthe linear acceleration {umlaut over (x)}₂ of the flywheel 314 asrepresented by Equation 150.θ+c=rx ₂  (Equation 130){dot over (θ)}=r{dot over (x)} ₂  (Equation 140){umlaut over (θ)}=r{umlaut over (x)} ₂  (Equation 150)

A flywheel 314 to piston 216 compliance force F₃ may be computed usingEquation 160 as the product of the flywheel rotational stiffness Z andthe difference between flywheel translation x₂ and piston translationx₁. For the example apparatus of FIG. 22 wherein the inerter (e.g., theflywheel 314) is integrated into the actuator 202, the flywheel 314moves with the piston 216 such that the flywheel translation x₂ and thepiston translation x₁ are the same, as indicated below in Equation 190.In this regard, the piston compliance force F₃ is zero (0) due to theassumption that x₂=x₁ as indicated below in Equation 190.F ₃ =Z(x ₂ −x ₁)  (Equation 160)

Substituting Equations 140 and 150 for flywheel velocity {dot over (θ)}and flywheel acceleration {umlaut over (θ)} into Equation 120, theflywheel reacted force F₂ may be expressed as follows:F ₂ =r ²(J{umlaut over (x)} ₂ +B{dot over (x)} ₂)  (Equation 170)

The piston reacted force F₁ may be computed as the sum of the product ofthe actuator (e.g., the piston) reacted inertia M at the rod end 214 andthe piston acceleration {umlaut over (x)}₁, the product of the actuator(e.g., the piston) resisting force C and the piston velocity {dot over(x)}, and the product of the actuator stiffness K and the pistondisplacement x₁, as shown in Equation 180:F ₁ =M{umlaut over (x)} ₁ +C{dot over (x)} ₁ +Kx ₁  (Equation 180)

As mentioned above, for the example shown in FIG. 22 wherein the inerter(e.g., the flywheel 314 and threaded shaft 322) is integrated into theactuator 202 such that the flywheel 314 and the piston 216 move inunison, the flywheel translation x₂ and the piston translation x₁ arethe same as represented by Equation 190. In addition the rod end 214 andthe piston 216 move in unison as represented by Equation 200. The capend 212 at x₀ is assumed to be fixed (e.g., non-translating) asrepresented by Equation 210.x ₂ =x ₁  (Equation 190){dot over (x)} ₁ ={dot over (x)}  (Equation 200){umlaut over (x)} ₀ ={dot over (x)} ₀ =x ₀=0  (Equation 210)

Performing a Laplace transform on a differential equation (not shown)representing the natural frequency of the example apparatus shown inFIG. 22, the resulting transfer function

$\frac{X(s)}{F(s)}$is expressed as shown in Equation 220 wherein X(s) represent theresponse of the apparatus of FIG. 22 and F(s) represents the input tothe apparatus:

$\begin{matrix}{\frac{X(s)}{F(s)} = \frac{\frac{1}{{r^{2}J} + M}}{s^{2} + {\frac{{r^{2}B} + C}{{r^{2}J} + M}s} + \frac{K}{{r^{2}J} + M}}} & \left( {{Equation}\mspace{14mu} 220} \right)\end{matrix}$

The natural frequency ω_(n) of oscillation of the example apparatus ofFIG. 22 may be expressed as shown in Equation 230 wherein K is theactuator stiffness, r is the thread rate, and J is the flywheelrotational inertia, as described above.

$\begin{matrix}{\omega_{n} = \left( \frac{K}{{r^{2}J} + M} \right)^{1/2}} & \left( {{Equation}\mspace{14mu} 230} \right)\end{matrix}$

Equation 240 represents the damping factor ζ of the example apparatus ofFIG. 22 which characterizes the decay in oscillatory response to theinput (e.g., flutter of a flight control surface).

$\begin{matrix}{\zeta = \frac{{r^{2}B} + C}{2\left( {K\left( {{r^{2}J} + M} \right)} \right)^{1/2}}} & \left( {{Equation}\mspace{14mu} 240} \right)\end{matrix}$

FIG. 23 is a graph plotting frequency 380 vs. magnitude 382 (amplitude)of the oscillatory response to a dynamic load for an actuator 202operating under three (3) different working pressures (3000 psi, 5000psi, and 8000 psi). The vertical centerline represents a flutterfrequency of 20 Hz corresponding to the dynamic load. The plots of FIG.23 illustrate the reduction in response amplitude 384 provided by theactuator 202 with integrated inerter 300 of FIG. 22, relative to theresponse amplitude for the same actuator operating without an inerter.The reduction in response amplitude represents an optimization based onsetting the response amplitude at the flutter frequency for the actuator202 operating at 8000 psi with an inerter 300 equal to the responseamplitude at the flutter frequency for the actuator 202 operating at3000 psi without the inerter 300, and optimizing the thread pitch r ofthe threaded shaft 322, the flywheel rotational inertia J, and thedamping factor ζ (Equation 240). For the actuator 202 operating at 8000psi, the inerter 300 facilitates a reduction in response amplitude 384of almost 5 dB at the flutter frequency of 20 Hz.

FIG. 24 is a flowchart having one or more operations that may beincluded in a method 400 of damping an actuator 202 using an inerter300. As mentioned above, the damping of the actuator 202 may comprisereducing actuator load oscillatory amplitude using inerter 300. Asindicated above, in some examples, the inerter 300 may be a separatecomponent from the actuator 202 and coupled to the same movable device124 as the actuator 202 (e.g., FIGS. 1 and 4-9). In other examples, theinerter 300 may be integrated into the actuator 202 (e.g., FIGS. 2 and10-22).

Step 402 of the method 400 includes actuating the movable device 124using an actuator 202. In the example of a flight control system 120 ofan aircraft 100, the method may include using a linear actuator such asa linear hydraulic actuator 204 or a linear electro-mechanical actuator242. For example, FIGS. 4-6 illustrate a linear hydraulic actuator 204configured to actuate an aileron 130 pivotably mounted to a wing 114 ofan aircraft 100. However, as mentioned above, the movable device 124 maybe any type of movable device that may be actuated by an actuator 202.

Step 404 of the method 400 includes axially accelerating, using aninerter 300 coupled to the movable device 124, the first terminal 302 ofthe inerter 300 relative to the second terminal 304 of the inerter 300.As indicated above, the inerter 300 may be coupled between the supportstructure 116 and the movable device 124 (e.g., FIGS. 4 and 6). Forexample, the first terminal 302 may be coupled to the movable device 124and the second terminal 304 may be coupled to the support structure 116,or the first terminal 302 may be coupled to the support structure 116and the second terminal 304 may be coupled to the movable device 124.Alternatively, the inerter 300 may be integrated into the actuator 202(e.g., FIGS. 10-21) which may be coupled between the support structure116 and the movable device 124. In such examples, as mentioned above,the rod end 214 or the cap end 212 of the actuator 202 functions as(e.g., is one and the same as) the first terminal 302 of the inerter300, and the remaining rod end 214 or cap end 212 of the actuator 202functions as (e.g., is one and the same as) the second terminal 304 ofthe inerter 300.

Step 406 of the method 400 includes rotationally accelerating theflywheel 314 simultaneous with the axial acceleration of the firstterminal 302 relative to the second terminal 304. Because the inerter300 and the actuator 202 are coupled to the same movable device 124(e.g., FIGS. 1 and 4-9) or because the inerter 300 is integrated intothe actuator 202 (e.g., FIGS. 2 and 10-21), the axial acceleration ofthe first terminal 302 relative to the second terminal 304 issimultaneous with and in proportion to the actuation of the movabledevice 124 by the actuator 202. In this regard, the flywheel 314rotationally accelerates and decelerates in proportion to the axialacceleration and deceleration of the first terminal 302 relative to thesecond terminal 304 in correspondence with the actuation of the movabledevice 124 by the actuator 202.

Step 408 of the method 400 includes damping the movement of the actuator202 in response to rotating the flywheel 314. In one example, the methodmay include reducing actuator load oscillatory amplitude of the movabledevice 124 in response to rotationally accelerating the flywheel 314.Regardless of whether the inerter 300 is a separate component from theactuator 202 or the inerter 300 is integrated into the actuator 202, themethod may include rotationally accelerating the flywheel 314 in amanner reducing actuator load oscillatory amplitude at resonance of themovable device 124 coupled to the actuator 202. In one example, themethod may include reducing actuator load oscillatory amplitude by atleast 50% relative to the oscillatory amplitude for the movable device124 actuated by the same actuator but without the inerter, as mentionedabove. The inerter 300 may be configured to reduce actuator loadoscillatory amplitude at a resonant frequency of up to approximately 20Hz (e.g., ±5 Hz). The movable device 124 may be a flight control surface122 (e.g., a hydraulically-actuated aileron 130) of an aircraft 100 andthe resonance (e.g., the resonant frequency) may correspond to flutterof the flight control surface 122 as induced by aerodynamic forcesacting on the flight control surface 122.

As mentioned above, in examples where the inerter 300 is integrated intothe actuator 202, the flywheel 314 may include a plurality of flywheelprotrusions 320 (e.g., flywheel blades—see FIGS. 11-12) extendingoutwardly from the flywheel 314. The flywheel 314 and the flywheelprotrusions 320 may be immersed in hydraulic fluid contained within thecap end chamber 236. In such examples, the method may include rotatingthe flywheel 314 within the hydraulic fluid and generating or increasingviscous damping of the actuator 202 movement in response to rotating theflywheel 314 in correspondence with the actuation of the movable device124. The viscous damping may contribute toward the damping provided bythe rotational inertia of the flywheel 314.

In still other examples, the method may include actively controlling therotation of the flywheel 314 in correspondence with relative axialmovement of the piston rod 224 and threaded shaft 322. For example, theinerter 300 may include or incorporate an electric flywheel motor 350 asdescribed above in the examples illustrated in FIGS. 15-16 and 19-20. Insome examples, as mentioned above, the actuator 202 may include a linearposition sensor (not shown) configured to sense the linear position ofthe piston 216 within the actuator 202 and generate a signalrepresentative of the piston position. The method may includecommutating the flywheel motor 350 in correspondence with the linearposition of the piston 216 as represented by the signal generated by theposition sensor.

Active control of the flywheel 314 rotation may include acceleratingand/or decelerating the flywheel 314 using the flywheel motor 350. Forexample, the flywheel motor 350 may be operated in a manner to apply atorque to the flywheel 314 in correspondence with or in the direction ofrotation of the flywheel 314. In this regard, the flywheel motor 350 mayassist a commanded direction of motion of the actuator 202. In someexamples, active control of flywheel rotation may include acceleratingthe flywheel 314 during initiation of actuation by the actuator 202 ofthe movable device 124 toward a commanded position. In this regard, theflywheel motor 350 may rotationally accelerate the flywheel 314 at thestart of axial acceleration of the first terminal 302 relative to secondterminal 304 by an amount at least partially or completely eliminatingthe force generated at the first terminal 302 and second terminal 304due to actuation of the movable device 124 by the actuator 202. By usingthe flywheel motor 350 to rotationally accelerate the flywheel 314 atthe start of axial acceleration, the force required to axially move thefirst terminal 302 relative to the second terminal 304 may be reduced oreliminated which may increase the speed at which the actuator 202 movesthe movable device 124 toward a commanded position.

Alternatively, the flywheel motor 350 may be operated in a manner toapply a torque to the flywheel 314 in a direction opposite the rotationof the flywheel 314. In this regard, the application of motor-generatedtorque in a direction opposite the rotation of the flywheel 314 mayresist the torque generated by the relative axial acceleration of thefirst terminal 302 and second terminal 304. In this regard, activecontrol by the flywheel motor 350 may oppose the terminal-developedtorque at the end of actuator 202 motion when the commanded position isreached. In this manner, the step of actively controlling rotation ofthe flywheel 314 may include using the flywheel motor 350 to dynamicallybrake or decelerate the flywheel 314 as the actuator 202 approaches acommanded position to prevent position overshoot.

In a further example, active control of flywheel 314 rotation mayinclude using a brake 360 (e.g., FIGS. 16 and 20) to decelerate theflywheel 314 as the actuator 202 approaches a commanded position of themovable device 124 to prevent position overshoot of the commandedposition. The method may additionally include dynamically braking therotation of the flywheel 314 such as to oppose disturbances (e.g.,undesirable motion) of the actuator 202. The step of dynamically braking(e.g., decelerating or reducing rotational speed) of the flywheel 314may be performed using a brake 360 operatively engageable to theflywheel 314 (e.g., FIGS. 16 and 20) or operatively engageable to abrake rotor (not shown) that may be fixedly coupled to the flywheel 314.Alternatively or additionally, the step of dynamically braking theflywheel 314 may be performed using rotational drag generated by theflywheel motor 350 as described above.

Now referring to FIG. 25, FIG. 25 is a perspective view of an aircraft100 having one or more translational inerter assemblies 500 (see FIGS.26-31) for damping movement 606 (see FIG. 31) of the flight controlsurface 122 of the aircraft 100. The aircraft 100 may include thefuselage 102 and the pair of wings 114 extending outwardly from thefuselage 102. The aircraft 100 may include a pair of propulsion units115 (e.g., gas turbine engines). As mentioned above, each wing 114 mayinclude one or more movable devices 124 configured as flight controlsurfaces 122 of the flight control system 120, where the flight controlsurfaces 122 may be actuated by one or more actuators 202 (see FIG. 26)that may be coupled to the one or more translational inerter assemblies500 (see FIGS. 26-31). The flight control surface 122 (see FIGS. 25, 26)may be hingedly coupled to a support structure 116 (see FIGS. 25, 26)that is rigid, such as the wing 114, including a wing portion supportstructure 117 (see FIG. 26), such as the wing spar 118 (see FIG. 26), orsuch as another suitable support structure.

Such flight control surfaces 122 (see FIG. 25) on the wings 114 (seeFIG. 25) may include, but are not limited to, spoilers, ailerons 130(see FIG. 25), and one or more high-lift devices, such as a leading edgeslats and/or trailing edge flaps. At the aft end of the fuselage 102(see FIG. 25), the empennage 104 (see FIG. 25) may include one or morehorizontal tails 110 (see FIG. 25) and the vertical tail 106 (see FIG.25), any one or more of which may include flight control surfaces 122(see FIG. 25) such as the elevator 112 (see FIG. 25), the rudder 108(see FIG. 25), or other types of movable devices 124 (see FIG. 25) thatmay be actuated by one or more actuators 202 (see FIG. 26) that may becoupled to one or more translational inerter assemblies 500 (see FIGS.26-31).

Now referring to FIG. 26, FIG. 26 is a top view of a wing section 114 aof the wing 114, taken along line 26-26 of FIG. 25, illustrating anactuator 202 and two translational inerter assemblies 500, such as inthe form of a first translational inerter assembly 500 a and a secondtranslational inerter assembly 500 b, operatively coupled to the flightcontrol surface 122 in the form of aileron 130. FIG. 26 shows the flightcontrol system 120 comprising the movable device 124, such as the flightcontrol surface 122 in the form of aileron 130. As shown in FIG. 26, theflight control surface 122 may be hingedly coupled to the supportstructure 116 such as the wing 114, including the wing portion supportstructure 117, such as the wing spar 118, or such as another suitablesupport structure. The flight control surface 122 (see FIG. 26) may bepivotable or rotatable about a hinge axis 126 (see FIGS. 27-30A). Theflight control surface 122 preferably comprises the aileron 130, and thesupport structure 116 preferably comprises the wing 114, the wingportion support structure 117, the wing spar 118, or another suitablesupport structure.

As shown in FIG. 26, the flight control surface 122, such as the aileron130, comprises a first end 502 a and a second end 502 b. As furthershown in FIG. 26, the first translational inerter assembly 500 a isattached to the first end 502 a of the flight control surface 122, suchas the aileron 130, and the second translational inerter assembly 500 bis attached to the second end 502 b of the flight control surface 122,such as the aileron 130. As further shown in FIG. 26, the translationalinerter assembly 500 comprises one or more attachment points 504 forconnecting the flight control surface 122 to the support structure 116,such as the wing 114, including the wing portion support structure 117,such as the wing spar 118, or such as another suitable supportstructure, and to provide additional structural robustness. For example,as shown in FIG. 26, the translational inerter assembly 500, such as inthe form of first translational inerter assembly 500 a, may compriseattachment point 504, such as in the form of a first bearing 504 a thatattaches the first end 502 a of the flight control surface 122 to thesupport structure 116. In addition, as further shown in FIG. 26 thetranslational inerter assembly 500, such as in the form of secondtranslational inerter assembly 500 b, may comprise attachment point 504,such as in the form of a second bearing 504 b that attaches the secondend 502 b of the flight control surface 122 to the support structure116. The attachment points 504 (see FIG. 26), such as in the form of thefirst bearing 504 a (see FIG. 26) and the second bearing 504 b (see FIG.26), constrain the flight control surface 122 in five (5) degrees offreedom and allow rotation about the hinge axis 126 (see FIGS. 27-30A).

As further shown in FIG. 26, the translational inerter assembly 500,such as in the form of first translational inerter assembly 500 a, maybe fixedly attached to the support structure 116 via a first torsion barbearing 506 a, and the translational inerter assembly 500, such as inthe form of second translational inerter assembly 500 b, may be fixedlyattached to the support structure 116 via a second torsion bar bearing506 b. The flight control surface 122 (see FIG. 26), such as in the formof aileron 130 (see FIG. 26), may be actuated by one or more actuators202 (see FIG. 26) located on one or both ends of the flight controlsurface 122 (see FIG. 26), such as the aileron 130 (see FIG. 26). InFIG. 26, the actuator 202 and the translational inerter assemblies 500are provided as separate components and may each be coupled between thesupport structure 116, such as the wing portion support structure 117,the wing spar 118, or another suitable support structure, and the flightcontrol surface 122, such as the aileron 130. The one or more actuators202 (see FIG. 26) are present, along with the translational inerterassembly 500. The one or more actuators 202 may be positioned anywherealong the wing spar 118 and the aileron 130 for desired and advantageousstructural strength, rigidity, aircraft weight, and/or installationcost. The translational inerter assembly 500 may be applied to anyflight control surface that comprises a flutter critical surface.

Now referring to FIG. 27, FIG. 27 is a sectional view of the wingsection 114 a of the wing 114, taken along line 27-27 of FIG. 26, andillustrates an example of the translational inerter assembly 500, suchas the first translational inerter assembly 500 a, installed in a firstend opening 508 a of the first end 502 a (see FIG. 26) of the movabledevice 124, such as the flight control surface 122, for example, theaileron 130. FIG. 27 shows the translational inerter assembly 500, suchas the first translational inerter assembly 500 a, comprising a pressfit element 510 inserted within the first end opening 508 a, an inertiaelement 512 inserted within the press fit element 510, and a torsion bar514, or torsion rod, inserted within the inertia element 512. FIG. 27further shows the attachment point 504, such as in the form of the firstbearing 504 a, that attaches the flight control surface 122, such as theaileron 130, to the support structure 116, such as the wing portionsupport structure 117 of the wing 114, to provide further structuralrobustness. FIG. 27 further shows the translational inerter assembly500, such as the first translational inerter assembly 500 a, attached toa horn element 516 and a bearing 518 that may be attached to one or moreactuators 202 (see FIGS. 26, 28) for surface position control.

FIG. 28 is a sectional view of the wing section 114 a of the wing 114,taken along line 28-28 of FIG. 26, and illustrates an example of thetranslational inerter assembly 500, such as the second translationalinerter assembly 500 b, installed in a second end opening 508 b of thesecond end 502 b (see FIG. 26) of the movable device 124, such as theflight control surface 122, for example, the aileron 130. FIG. 28 showsthe translational inerter assembly 500, such as the second translationalinerter assembly 500 b, comprising the press fit element 510 insertedwithin the second end opening 508 b, the inertia element 512 within thepress fit element 510, and the torsion bar 514 within the inertiaelement 512. FIG. 28 further shows the attachment point 504, such as inthe form of the second bearing 504 b, that attaches the flight controlsurface 122, such as the aileron 130, to the support structure 116, suchas the wing portion support structure 117 of the wing 114, to providefurther structural robustness.

FIG. 28 further shows an example of the actuator 202, such as in theform of a hydraulic actuator 204 that is linear, and that ismechanically coupled between the support structure 116, such as the wingspar 118, and one end of the movable device 124, such as the flightcontrol surface 122, for example, the aileron 130. As shown in FIG. 28,the actuator 202, such as in the form of hydraulic actuator 204,includes a piston 216 coupled to a piston rod 224. The piston 216 (seeFIG. 28) is slidable within an actuator housing 228 (see FIG. 7) (e.g.,a cylinder). The actuator 202 (see FIG. 28), such as in the form ofhydraulic actuator 204 (see FIG. 28), further includes a rod end 214(see FIG. 28) and a cap end 212 (see FIG. 28) axially movable relativeto one another in response to pressurized hydraulic fluid acting in anunbalanced manner on one or both sides of the piston 216 (see FIG. 28)inside the actuator housing 228 (see FIG. 7). In the example shown inFIG. 28, the rod end 214 of the actuator 202, such as the hydraulicactuator 204, is coupled to a bellcrank 128. The bellcrank 128 (see FIG.28) is hingedly coupled to the flight control surface 122 (see FIG. 28),such as the aileron 130 (see FIG. 28), in a manner such that linearactuation of the hydraulic actuator 204 (see FIG. 28) causes pivoting ofthe flight control surface 122, such as the aileron 130, about the hingeaxis 126 (see FIG. 28). Alternatively, instead of the bellcrank 128, thehorn element 516 (see FIG. 27) and bearing 518 (see FIG. 27) may be usedas a connection between the rod end 214 of the actuator 202 and thetranslational inerter assembly 500 and the flight control surface 122.The cap end 212 (see FIG. 28) of the actuator 202 (see FIG. 28), such asthe hydraulic actuator 204 (see FIG. 28), is coupled to the wing spar118 (see FIG. 28).

Now referring to FIG. 29A, FIG. 29A is a schematic perspective view ofthe movable device 124, such as the flight control surface 122, forexample, the aileron 130, with the translational inerter assembly 500,such as the first translational inerter assembly 500 a, installed in thefirst end opening 508 a of the first end 502 a, of the movable device124, such as the flight control surface 122, for example, the aileron130. FIG. 29A further shows the first translational inerter assembly 500a comprising the press fit element 510, in the form of a first press fitelement 510 a, inserted within the first end opening 508 a, comprisingthe inertia element 512, in the form of a first inertia element 512 a,inserted within the press fit element 510, comprising the torsion bar514 or torsion rod, in the form of a first torsion bar 514 a or firsttorsion rod, inserted within the inertia element 512 at one end andattached to the first torsion bar bearing 506 a at the other end. Thefirst torsion bar bearing 506 a (see FIG. 29A) is attached to thesupport structure 116 (see FIG. 29A). The first torsion bar bearing 506a (see FIG. 29A) constrains the torsion bar 514 in six (6) degrees offreedom. FIG. 29A further shows the first translational inerter assembly500 a comprising the splined attachment element 520, in the form of afirst splined attachment element 520 a, coupled to the torsion bar 514.

As shown in FIG. 29A, the translational inerter assembly 500, such as inthe form of first translational inerter assembly 500 a, may furthercomprise attachment point 504, in the form of the first bearing 504 a,that attaches the first end 502 a of the flight control surface 122 tothe support structure 116. In addition, as further shown in FIG. 29A thetranslational inerter assembly 500, such as in the form of secondtranslational inerter assembly 500 b, may further comprise attachmentpoint 504, such as in the form of a second bearing 504 b, that attachesthe second end 502 b of the flight control surface 122 to the supportstructure 116. The attachment points 504 (see FIG. 29A), such as in theform of the first bearing 504 a (see FIG. 29A) and the second bearing504 b (see FIG. 29A), constrain the flight control surface 122 in five(5) degrees of freedom and allow rotation about the hinge axis 126 (seeFIG. 29A).

Now referring to FIG. 29B, FIG. 29B is a schematic perspective view ofthe movable device 124, such as the flight control surface 122, forexample, the aileron 130, of FIG. 29A, with two translational inerterassemblies 500, such as in the form of the first translational inerterassembly 500 a, installed in the first end opening 508 a of the firstend 502 a, and such as in the form of the second translational inerterassembly 500 b, installed in the second end opening 508 b of the secondend 502 b, of the flight control surface 122. Similar to FIG. 29A, FIG.29B shows the first translational inerter assembly 500 a comprising thepress fit element 510, in the form of first press fit element 510 a,inserted within the first end opening 508 a, comprising the inertiaelement 512, in the form of first inertia element 512 a, inserted withinthe press fit element 510, and comprising the torsion bar 514, in theform of first torsion bar 514 a, inserted within the inertia element 512at one end and attached to the first torsion bar bearing 506 a at theother end. The first torsion bar bearing 506 a (see FIG. 29B) isattached to the support structure 116 (see FIG. 29B). FIG. 29B furthershows the first translational inerter assembly 500 a comprising thesplined attachment element 520, in the form of first splined attachmentelement 520 a, coupled to the torsion bar 514. As shown in FIG. 29B, thefirst translational inerter assembly 500 a, may further compriseattachment point 504, in the form of the first bearing 504 a, thatattaches the first end 502 a of the flight control surface 122 to thesupport structure 116.

FIG. 29B further shows the second translational inerter assembly 500 bcomprising the press fit element 510, in the form of a second press fitelement 510 b, inserted within the second end opening 508 b, comprisingthe inertia element 512, in the form of a second inertia element 512 b,inserted within the press fit element 510, and comprising the torsionbar 514 or torsion rod, in the form of a second torsion bar 514 b or asecond torsion rod, inserted within the inertia element 512 at one endand attached to a second torsion bar bearing 506 b at the other end. Thesecond torsion bar bearing 506 b (see FIG. 29B) is attached to thesupport structure 116 (see FIG. 29B). FIG. 29B further shows the secondtranslational inerter assembly 500 b comprising the splined attachmentelement 520, in the form of a second splined attachment element 520 b,coupled to the torsion bar 514. As further shown in FIG. 29B the secondtranslational inerter assembly 500 b may further comprise attachmentpoint 504, such as in the form of a second bearing 504 b, that attachesthe second end 502 b of the flight control surface 122 to the supportstructure 116. The attachment points 504 (see FIG. 29B), such as in theform of the first bearing 504 a (see FIG. 29B) and the second bearing504 b (see FIG. 29B), constrain the flight control surface 122 in five(5) degrees of freedom and allow rotation about the hinge axis 126 (seeFIG. 29B).

Now referring to FIG. 30A, FIG. 30A is an exploded perspective view ofthe movable device 124, such as the flight control surface 122, forexample, the aileron 130 a, with the translational inerter assembly 500,such as in the form of first translational inerter assembly 500 a, in adisassembled position 600. As shown in FIG. 30A, the movable device 124,such as the flight control surface 122, for example, the aileron 130 a,comprises the first end 502 a with the first end opening 508 a andcomprises the second end 502 b with the second end opening 508 b. Theflight control surface 122 (see FIG. 30A) comprises the hinge axis 126(see FIG. 30A). The translational inerter assembly 500 further comprisesthe first bearing 504 a (see FIG. 30A) attaching the first end 502 a(see FIG. 30A) of the flight control surface 122 (see FIG. 30A) to thesupport structure 116 (see FIG. 30B), and the second bearing 504 b (seeFIG. 30A) attaching the second end 502 b (see FIG. 30A) of the flightcontrol surface 122 to the support structure 116 (see FIG. 30B). Thefirst bearing 504 a (see FIGS. 30A-30B) and the second bearing 504 b(see FIG. 30A-30B) constrain the flight control surface 122 in five (5)degrees of freedom and allow rotation about the hinge axis 126 (see FIG.30A).

As further shown in FIG. 30A, the translational inerter assembly 500,such as in the form of first translational inerter assembly 500 a,comprises the press fit element 510, in the form of first press fitelement 510 a. As shown in FIG. 30A, the press fit element 510 comprisesa press fit element first end 522 a, a press fit element second end 522b, and a press fit element body 528 formed therebetween. The press fitelement first end 522 a (see FIG. 30A) has an outer face 524 a (see FIG.30A) with a plurality of fastener holes 526 (see FIG. 30A), for example,two fastener holes 526, configured to attach to, via a plurality offasteners 527 (see FIG. 30A), the horn element 516 (see FIG. 30A) andthe bearing 518 (see FIG. 30A) for surface position control using one ormore actuators 202 (see FIG. 28). The press fit element second end 522 b(see FIG. 30A) has an inner face 524 b (see FIG. 30A). As further shownin FIG. 30A, the press fit element 510 has an exterior 530, a helicalspline interior 532, and a press fit element through opening 534 havinga plurality of interior helical splines 536 formed along the press fitelement through opening 534. The press fit element 510 (see FIG. 30A)has a press fit element outer diameter 538 (see FIG. 30A) and a pressfit element length 540 (see FIG. 30A).

As further shown in FIG. 30A, the translational inerter assembly 500,such as in the form of first translational inerter assembly 500 a,comprises the inertia element 512, such as the first inertia element 512a. As shown in FIG. 30A, the inertia element 512 comprises an inertiaelement first end 542 a, an inertia element second end 542 b, and aninertia element body 544 formed therebetween. The inertia element body544 (see FIG. 30A) comprises a helical spline portion 546 (see FIG. 30A)integral with a shaft portion 548 (see FIG. 30A). The helical splineportion 546 has a plurality of exterior helical splines 550. Theplurality of exterior helical splines 550 may comprise one of, righthand direction helical splines 551 a, or left hand direction helicalsplines 551 b. If there are two translational inerter assemblies 500,such as in the form of first translational inerter assembly 500 a (seeFIG. 29B) and second translational inerter assembly 500 b (see FIG.29B), the plurality of exterior helical splines 550 of eachtranslational inerter assembly 500 are in opposite directions 553 (seeFIG. 31), that is one translational inerter assembly 500 has right handdirection helical splines 551 a and the other translational inerterassembly 500 has left hand direction helical splines. By having theinertia element 512 of one translational inerter assembly 500 with righthand direction helical splines 551 a and the inertia element 512 of theother translational inerter assembly 500 with left hand directionhelical splines, the inertia elements 512 move opposite one anotheralong the hinge axis 126 (see FIG. 30A) to keep the loads associatedwith the acceleration of the inertia elements 512 equal and opposite, sothat the loads nearly cancel to limit the loading of the first bearing504 a (see FIG. 30A) and the second bearing 504 b (see FIG. 30A).

The plurality of exterior helical splines 550 (see FIG. 30A) of theinertia element 512 (see FIG. 30A) correspond to the plurality ofinterior helical splines 536 (see FIG. 30A) of the press fit element 510(see FIG. 30A), when the inertia element 512 is inserted into andcoupled within the press fit element 510. As used herein, “helicalspline” means equally spaced grooves or projections that form a helixabout the exterior of a helical spine portion or a shaft portion of aninertia element, where the sides of the grooves or projections may beparallel or involute tooth form, and where the grooves or projectionsfit into corresponding grooves or projections in a helical splineinterior of a press fit element, enabling the exterior helical splinesof the inertia element to translate or slide along a plurality ofinterior helical splines of the helical spline interior of the press fitelement.

As further shown in FIG. 30A, the inertia element 512 has an exterior552, a linear spline interior 554, and an inertia element throughopening 556 having a plurality of interior linear splines 558 formedalong the inertia element through opening 556. The inertia element 512(see FIG. 30A) has an inertia element inner diameter 560 (see FIG. 30A)and an inertia element length 562 (see FIG. 30A). The inertia elementinner diameter 560 is smaller than the press fit element outer diameter538, and the inertia element length 562 is shorter than the press fitelement length 540.

As further shown in FIG. 30A, the translational inerter assembly 500,such as in the form of first translational inerter assembly 500 a,comprises the torsion bar 514, such as the first torsion bar 514 a. Asshown in FIG. 30A, the torsion bar 514 comprises a torsion bar first end564 a with a first end face 566 a, a torsion bar second end 564 b with asecond end face 566 b, and a torsion bar body 568 formed therebetween.The torsion bar body 568 (see FIG. 30A) comprises a linear splineexterior 570 (see FIG. 30A and a solid interior 574 (see FIG. 30A). Thelinear spline exterior 570 comprises a plurality of exterior linearsplines 572 longitudinally formed along a torsion bar length 578 of thetorsion bar 514. The torsion bar 514 (see FIG. 30A) has a torsion barouter diameter 576 (see FIG. 30A). The torsion bar outer diameter 576 issmaller than the inertia element inner diameter 560, and the torsion barlength 578 is longer than the inertia element length 562.

Preferably, there is a balance of surface areas between, (a) the contactsurface area or surface interface between the plurality of exteriorhelical splines 550 of the inertia element 512 and the plurality ofinterior helical splines 536 of the press fit element 510, and (b) thecontact surface area or surface interface between the interior linearsplines 558 of the inertia element 512 and the plurality of exteriorlinear splines 572 of the torsion bar 514, such that the balance resultsin a wear rate that is equivalent, or substantially equivalent, of boththe interior and exterior spline surface interfaces of the inertiaelement 512. The balance of surface areas may be achieved in the contactsurface area or surface interface between the interior splines and thecontact surface area or surface interface between the exterior splines,given the larger radius of the exterior helical splines 550 of theinertia element 512 and a shorter length spline engagement to assure anappropriate balance of surface area. The surface area balance maypreferably be that which achieves a maximum life and acceptable load onthe shaft portion 548 or interior spline shaft of the inertia element512 and on the exterior linear splines 572 of the torsion bar 514. Atorsion bar outer diameter 576 (see FIG. 30A) may be reduced near thesplined attachment element 520 (see FIG. 30A) inner diameter.

As further shown in FIG. 30A, the translational inerter assembly 500,such as in the form of first translational inerter assembly 500 a,comprises the splined attachment element 520, such as the first splinedattachment element 520 a. As shown in FIG. 30A, the splined attachmentelement 520 comprises a splined attachment element first end 580 a, asplined attachment element second end 580 b, and a splined attachmentelement body 582 formed therebetween. The splined attachment elementbody 582 (see FIG. 30A) comprises a fastener portion 584 (see FIG. 30A)integral with a shaft portion 588 (see FIG. 30A). The fastener portion584 has a plurality of fastener holes 586 (see FIG. 30A), for example,four fastener holes 586, configured to attach to, via a plurality offasteners 587 (see FIG. 30A), the first torsion bar bearing 506 a (seeFIG. 30A) and/or the support structure 116 (see FIG. 30B). Although FIG.30A shows only one fastener 587, each fastener hole 586 would beconfigured to receive a separate fastener 587.

As further shown in FIG. 30A, the splined attachment element 520 has anexterior 590, a linear spline interior 592, and a splined attachmentelement through opening 594 having a plurality of interior linearsplines 595 formed along the splined attachment element through opening594. The splined attachment element 520 (see FIG. 30A) has a splinedattachment element inner diameter 596 (see FIG. 30A) and a splinedattachment element length 598 (see FIG. 30A). The splined attachmentelement inner diameter 596 is larger than the torsion bar outer diameter576, and the splined attachment element length 598 is shorter than thetorsion bar length 578.

Rotation of the flight control surface 122 (see FIG. 30B) causestranslational movement 604 (see FIG. 31) of the inertia element 512 (seeFIG. 30B), via the press fit element 510 (see FIG. 30B), along the hingeaxis 126 (see FIG. 30B) of the flight control surface 122, and along thetorsion bar 514 (see FIG. 30B). This results in the translationalinerter assembly 500 (see FIG. 30B) damping movement 606 (see FIG. 31)of the flight control surface 122 (see FIG. 30B).

Now referring to FIG. 30B, FIG. 30B is a perspective view of the movabledevice 124, such as the flight control surface 122, for example, theaileron 130 a, with the translational inerter assembly 500, such as inthe form of first translational inerter assembly 500 a, of FIG. 30A,with the translational inerter assembly in an assembled position 602.FIG. 30B shows the movable device 124, such as the flight controlsurface 122, for example, the aileron 130 a, comprising the first end502 a with the first end opening 508 a and comprising the second end 502b with the second end opening 508 b. The first end 502 a (see FIG. 30B)of the flight control surface 122 (see FIG. 30B) is attached toattachment point 504 (see FIG. 30B), such as in the form of firstbearing 504 a (see FIG. 30B), and the first bearing 504 a is attached tothe support structure 116 (see FIG. 30B), such as the wing 114 (see FIG.30B), for example, the wing portion support structure 117 (see FIG.30B). The second end 502 b (see FIG. 30B) of the flight control surface122 (see FIG. 30B) is attached to attachment point 504 (see FIG. 30B),such as in the form of second bearing 504 b (see FIG. 30B), and thesecond bearing 504 b is attached to the support structure 116 (see FIG.30B), such as the wing 114 (see FIG. 30B), for example, the wing portionsupport structure 117 (see FIG. 30B).

As further shown in FIG. 30B, the press fit element 510, such as thefirst press fit element 510 a, of the translational inerter assembly500, such as in the form of first translational inerter assembly 500 a,is inserted into and press fit within the first end opening 508 a of theflight control surface 122. The press fit element 510 (see FIG. 30B) isfixedly disposed within the first end opening 508 a (see FIG. 30B) ofthe first end 502 a (see FIG. 30B) of the flight control surface 122 androtatably movable with the flight control surface 122.

In one version, as shown in FIGS. 29A and 30B, the press fit element 510may be inserted into one end of the flight control surface 122. Inanother version, as shown in FIG. 29B, the press fit element 510 may beinserted into both ends of the flight control surface 122. FIG. 30Bshows the press fit element first end 522 a and the press fit elementsecond end 522 b. The outer face 524 a (see FIG. 30A) of the press fitelement first end 522 a (see FIGS. 30A-30B) has the plurality offastener holes 526 (see FIG. 30B), for example, two fastener holes 526,configured to attach to, via the plurality of fasteners 527 (see FIG.30B), the horn element 516 (see FIG. 30B) and the bearing 518 (see FIG.30B) for surface position control using one or more actuators 202 (seeFIG. 28). The press fit element 510 (see FIG. 30B) has the plurality ofinterior helical splines 536 (see FIG. 30B).

As further shown in FIG. 30B, the inertia element 512, such as the firstinertia element 512 a, of the translational inerter assembly 500, suchas in the form of first translational inerter assembly 500 a, isinserted within the press fit element through opening 534 (see FIG. 30A)of the press fit element 510. The inertia element 512 is coupled to andinstalled in the press fit element 510. FIG. 30B shows the inertiaelement first end 542 a, the inertia element second end 542 b, and theinertia element through opening 556. The plurality of exterior helicalsplines 550 (see FIG. 30B) of the inertia element 512 (see FIG. 30B)correspond to and align with the plurality of interior helical splines536 (see FIG. 30B) of the press fit element 510 (see FIG. 30B) when theinertia element 512 is inserted within the press fit element 510. Asdiscussed above, the plurality of exterior helical splines 550 (see FIG.30B) may comprise one of, right hand direction helical splines 551 a(see FIG. 31), or left hand direction helical splines 551 b (see FIG.31).

As further shown in FIG. 30B, the torsion bar 514, such as the firsttorsion bar 514 a, of the translational inerter assembly 500, such as inthe form of first translational inerter assembly 500 a, has the torsionbar first end 564 a exteriorly positioned outside the inertia element512, the press fit element 510, and the first end opening 508 a of theflight control surface 122. The torsion bar 514 (see FIG. 30B) has thetorsion bar first end 564 a (see FIG. 30B) coupled to and installed inthe inertia element 512 (see FIG. 30B). As further shown in FIG. 30B,the torsion bar 514 has the torsion bar second end 564 b insertedthrough the inertia element through opening 556 and positioned withinthe inertia element 512, within the press fit element 510, and withinthe first end opening 508 a of the flight control surface 122. Thetorsion bar second end 564 b (see FIGS. 30A-30B) is also preferablyfixedly attached to the support structure 116 (see FIGS. 25, 30B) of anaircraft 100 (see FIG. 25). The torsion bar second end 564 b (see FIG.30B) is preferably fixedly attached to the support structure 116 (seeFIGS. 25, 30B) of the aircraft 100 (see FIG. 25), via the first torsionbar bearing 506 a (see FIG. 30B) attached to the splined attachmentelement 520 (see FIG. 30B). The first torsion bar bearing 506 a (seeFIG. 30B) constrains the torsion bar 514 (see FIG. 30B) in six (6)degrees of freedom.

The plurality of exterior linear splines 572 (see FIG. 30B)longitudinally formed along the linear spline exterior 570 (see FIG.30A) of the torsion bar 514 (see FIGS. 30A-30B) correspond to and alignwith the plurality of interior linear splines 558 (see FIG. 30A) of theinertia element 512 (see FIGS. 30A-30B) when the torsion bar 514 ispartially inserted within the inertia element 512. The plurality ofexterior linear splines 572 (see FIG. 30B) of the torsion bar 514 (seeFIG. 30B) and the plurality of interior linear splines 558 (see FIG.30A) of the inertia element 512 (see FIG. 30B) permit or allow atranslational movement 604 (see FIG. 31) of the inertia element 512along the torsion bar 514 (see FIG. 30B), but restrict rotation to thatpermitted by a torsion bar flexure 573 (see FIG. 31).

As further shown in FIG. 30B, the splined attachment element 520, suchas the first splined attachment element 520 a, of the translationalinerter assembly 500, such as in the form of first translational inerterassembly 500 a, is coupled to and around the torsion bar 514 near or atthe torsion bar first end 564 a. FIG. 30B shows the splined attachmentelement first end 580 a, the splined attachment element second end 580b, the fastener portion 584 integral with the shaft portion 588, and thesplined attachment element through opening 594. The fastener portion 584(see FIG. 3B) has the plurality of fastener holes 586 (see FIG. 30A),for example, four fastener holes 586, configured to attach to, via theplurality of fasteners 587 (see FIG. 30B), the first torsion bar bearing506 a (see FIG. 30B) and/or the support structure 116 (see FIG. 30B).Each fastener hole 586 (see FIG. 30B) is preferably configured toreceive a separate fastener 587 (see FIG. 30B). The plurality ofexterior linear splines 572 (see FIG. 30B) longitudinally formed alongthe linear spline exterior 570 (see FIG. 30A) of the torsion bar 514(see FIGS. 30A-30B) correspond to and align with the plurality ofinterior linear splines 595 (see FIG. 30A) of the splined attachmentelement 520 (see FIGS. 30A-30B), when the torsion bar 514 is insertedthrough the splined attachment element through opening 594. In otherwords, the plurality of interior linear splines 595 (see FIG. 30A)correspond to the plurality of exterior linear splines 572 (see FIGS.30A-30B) of the torsion bar 514 (see FIGS. 30A-30B).

Now referring to FIG. 31, FIG. 31 is a block diagram of a flight controlsystem 120 of an aircraft 100, including one or more actuators 202configured to actuate, or actuating, a movable device 124, such as aflight control surface 122, for example, an aileron 130, about a hingeaxis 126, and further including a translational inerter assembly (TIA)500 for damping movement 606 of the movable device 124, such as theflight control surface 122, for example, the aileron 130. The aircraft100 (see FIG. 31) comprises the flight control surface 122 (see FIG. 31)pivotably coupled to a support structure 116 (see FIG. 31). As shown inFIG. 31, the support structure 116 may comprise a wing 114, a wingportion support structure 117, a wing spar 118, or another suitablesupport structure 116. As further shown in FIG. 31, the actuator 202comprises the cap end 212, the piston 216, the piston rod, 224, and therod end 214.

As shown in FIG. 31, the aircraft 100 further comprises at least onetranslational inerter assembly 500 for damping movement 606 of theflight control surface 122. The at least one translational inerterassembly 500, as discussed above, may comprise the first translationalinerter assembly 500 a attached at the first end opening 508 a of thefirst end 502 a of the flight control surface 122, such as the aileron130. The at least one translational inerter assembly 500, as discussedabove, may further comprise the second translational inerter assembly500 b attached at the second end opening 508 b of the second end 502 bof the flight control surface 122, such as the aileron 130. In oneversion the flight control system 120 (see FIG. 31) comprises onetranslational inerter assembly 500 attached to one end of the flightcontrol surface 122. In another version, the flight control system 120(see FIG. 31) comprises two translational inerter assemblies 500 eachattached to opposite ends of the flight control surface 122. Forexample, the two translational inerter assemblies 500 may comprise thefirst translational inerter assembly 500 a attached to the first end 502a of the flight control surface 122, and may comprise the secondtranslational inerter assembly 500 b attached to the second end 502 b ofthe flight control surface 122.

As shown in FIG. 31, the translational inerter assembly 500 of theaircraft 100 comprises attachment points 504, including the firstbearing 504 a attaching the first end 502 a of the flight controlsurface 122 to the support structure 116, and the second bearing 504 battaching the second end 502 b of the flight control surface 122 to thesupport structure 116. The first bearing 504 a (see FIG. 31) and thesecond bearing 504 b (see FIG. 31) constrain the flight control surface122 (see FIG. 31) in five (5) degrees of freedom and allow rotationabout the hinge axis 126 (see FIG. 31). In addition, the translationalinerter assembly 500 (see FIGS. 27, 31) may be attached to a hornelement 516 (see FIGS. 27, 31) and a bearing 518 (see FIGS. 27, 31) thatmay be attached to one or more actuators 202 (see FIGS. 26, 28, 31) forsurface position control.

As shown in FIG. 31, the translational inerter assembly 500 of theaircraft 100 further comprises the press fit element (PFE) 510 that ispreferably fixedly disposed within the first end 502 a of the flightcontrol surface 122 and rotatably movable with the flight controlsurface 122. The press fit element (PFE) 510 (see FIG. 31) may comprisea first press fit element (PFE) 510 a (see FIGS. 29B, 31) fixedlydisposed within the first end 502 a (see FIGS. 29B, 31) of the flightcontrol surface 122 (see FIGS. 29B, 31), and a second press fit element(PFE) 510 b (see FIGS. 29B, 31) fixedly disposed within the second end502 b (see FIGS. 29B, 31) of the flight control surface 122 (see FIGS.29B, 31). Each press fit element 510 comprises a press fit element (PFE)through opening 534 (see FIGS. 30A, 31) with a plurality of interiorhelical splines 536 (see FIGS. 30A, 31).

As shown in FIG. 31, the translational inerter assembly 500 of theaircraft 100 further comprises the inertia element (IE) 512 coupled toand installed in the press fit element 510. The inertia element (IE) 512(see FIG. 31) may comprise a first inertia element (IE) 512 a (see FIGS.29B, 31) translationally disposed within the press fit element (PFE)through opening 534 (see FIGS. 30A, 31) of the first press fit element510 a (see FIGS. 29B, 31) of the first translational inerter assembly500 a (see FIG. 31). The inertia element (IE) 512 (see FIG. 31) mayfurther comprise a second inertia element (IE) 512 b (see FIGS. 29B, 31)translationally disposed within the press fit element (PFE) throughopening 534 (see FIGS. 30A, 31) of the second press fit element 510 b(see FIGS. 29B, 31) of the second translational inerter assembly 500 b(see FIG. 31).

As shown in FIG. 31, each inertia element 512 has a helical splineportion 546 with a plurality of exterior helical splines 550 that areconfigured to correspond to and correspond to the plurality of interiorhelical splines 536 of the press fit element 510. The plurality ofexterior helical splines 550 (see FIG. 31) may comprise right handdirection helical splines 551 a (see FIG. 31), or may comprise left handdirection helical splines 551 b (see FIG. 31). In one version, the firstinertia element 512 a (see FIG. 31) may have the plurality of exteriorhelical splines 550 (see FIG. 31) comprising right hand directionhelical splines 551 a (see FIG. 31), and the second inertia element 512b (see FIG. 31) may have the plurality of exterior helical splines 550comprising left hand direction helical splines 551 b (see FIG. 31). Ifthe flight control surface 122 (see FIGS. 29B, 31) has two translationalinerter assemblies 500 (see FIGS. 29B, 31), for example, the firsttranslational inerter assembly 500 a (see FIGS. 29B, 31) with the firstinertia element 512 a (see FIGS. 29B, 31) and the second translationalinerter assembly 500 b (see FIGS. 29B, 31) with the second inertiaelement 512 b (see FIGS. 29B, 31), then the plurality of exteriorhelical splines 550 (see FIGS. 29B, 31) of the first inertia element 512a and the second inertia element 512 b are preferably oriented or formedin opposite directions 553 (see FIG. 31).

As shown in FIG. 31, each inertia element 512 further has a shaftportion 548, where the shaft portion 548 is preferably integral with thehelical spline portion 546. Alternatively, the shaft portion 548 (seeFIG. 31) and the helical spline portion 546 (see FIG. 31) may beseparate pieces or parts that are attached or coupled together via asuitable attachment means. As shown in FIG. 31, each inertia elementfurther has an inertia element (IE) through opening 556 with a pluralityof interior linear splines 558.

As shown in FIG. 31, the translational inerter assembly 500 of theaircraft 100 further comprises the torsion bar (TB) 514 having thetorsion bar first end 564 a (see FIG. 30B) coupled to and installed inthe inertia element (IE) through opening 556 (see also FIG. 30B) of theinertia element 512 (see also FIG. 30B), and having the torsion barsecond end 564 b (see FIG. 30B) fixedly attached to the supportstructure 116 (see also FIG. 30B) of the aircraft 100. As shown in FIG.31, the torsion bar (TB) 514 may comprise a first torsion bar (TB) 514 acoupled or attached to a first torsion bar (TB) bearing 506 a, and maycomprise a second torsion bar (TB) 514 b coupled or attached to a secondtorsion bar (TB) bearing 506 b. The first torsion bar bearing 506 a (seeFIG. 31) constrains the torsion bar 514 (see FIG. 31), such as in theform of first torsion bar 514 a (see FIG. 31), in six (6) degrees offreedom. The second torsion bar bearing 506 b (see FIG. 31) constrainsthe torsion bar 514 (see FIG. 31), such as in the form of second torsionbar 514 b (see FIG. 31), in six (6) degrees of freedom.

Each torsion bar 514 (see FIG. 31) has a plurality of exterior linearsplines 572 (see FIG. 31) configured to correspond to and correspondingto the plurality of interior linear splines 558 (see FIG. 31) of theinertia element 512 (see FIG. 31). The plurality of exterior linearsplines 572 (see FIGS. 30B, 31) longitudinally formed along the linearspline exterior 570 (see FIG. 30A) of the torsion bar 514 (see FIGS.30A-30B, 31) correspond to and align with the plurality of interiorlinear splines 558 (see FIGS. 30A, 31) of the inertia element 512 (seeFIGS. 30A-30B, 31), when the torsion bar 514 is partially insertedwithin the inertia element 512. The plurality of exterior linear splines572 (see FIG. 31) of the torsion bar 514 (see FIG. 31) and the pluralityof interior linear splines 558 (see FIG. 31) of the inertia element 512(see FIG. 30B) facilitate translational movement 604 (see FIG. 31) ofthe inertia element 512 along the torsion bar 514 (see FIG. 30B), butrestrict rotation to that permitted by a torsion bar flexure 573 (seeFIG. 31).

As shown in FIG. 31, the translational inerter assembly 500 of theaircraft 100 may further comprise a splined attachment element (SAE) 520coupled to the torsion bar 514. The splined attachment element (SAE) 520may comprise a first splined attachment element (SAE) 520 a coupled tothe first torsion bar (TB) 514 a of the first translational inerterassembly 500 a, and may comprise a second splined attachment element(SAE) 520 b coupled to the second torsion bar (TB) 514 b of the secondtranslational inerter assembly 500 b. Each splined attachment element(SAE) 520 (see FIG. 31) comprises a fastener portion 584 (see FIG. 31)integral with a shaft portion 588 (see FIG. 31). Alternatively, thefastener portion 584 (see FIG. 31) and the shaft portion 588 (see FIG.31) may be separate pieces or parts that are attached or coupledtogether via a suitable attachment means. Each splined attachmentelement (SAE) 520 (see FIG. 31) may further comprise a splinedattachment element (SAE) through opening 594 (see FIG. 31) having aplurality of interior linear splines 595 (see FIG. 31) formed along thesplined attachment element (SAE) through opening 594.

Rotation of the flight control surface 122 (see FIG. 31) causestranslational movement 604 (see FIG. 31) of the inertia element 512 (seeFIG. 31), via the press fit element 510 (see FIG. 31), along the hingeaxis 126 (see FIG. 31) of the flight control surface 122 and along thetorsion bar 514 (see FIG. 31). This results in the at least onetranslational inerter assembly 500 (see FIG. 31) damping movement 606(see FIG. 31) of the flight control surface 122. Damping movement 606(see FIG. 31) of the flight control surface 122 (see FIG. 31) preferablyprovides increased flutter suppression 610 (see FIG. 31) of the flightcontrol surface 122. This preferably results in an improved hydraulicapplication stability 612 (see FIG. 31) and an increased efficientflight control actuation 614 (see FIG. 31). The translational inerterassembly 500 (see FIG. 31) of the aircraft 100 (see FIG. 31) provides apassive solution 616 (see FIG. 31), that is, the translational inerterassembly 500 uses a passive means to change the dynamics of the flightcontrol system 120 (see FIG. 31) instead of active control elements,such as one or more actuators 202 (see FIG. 31) and valve size ordiameter of hydraulic actuators 204 (see FIG. 28).

Now referring to FIG. 32, FIG. 32 is a flowchart having one or moreoperations that may be included in a method 650 for damping movement 606(see FIG. 31) of a flight control surface 122 (see FIG. 31) of anaircraft 100 (see FIGS. 25, 31). As shown in FIG. 32, the method 650comprises step 652 of installing at least one translational inerterassembly 500 at a first end 502 a of the flight control surface 122.

As discussed in detail above, the at least one translational inerterassembly 500 (see FIGS. 29A-31) comprises the press fit element 510 (seeFIGS. 29A-31) fixedly disposed within the first end 502 a (see FIGS.29A-30B) of the flight control surface 122 (see FIGS. 29A-31) androtatably movable with the flight control surface 122. The at least onetranslational inerter assembly 500 (see FIGS. 29A-31) further comprisesthe inertia element 512 (see FIGS. 29A-31) coupled to and installed inthe press fit element 510. The inertia element 512 (see FIGS. 29A-31)has the plurality of exterior helical splines 550 (see FIGS. 30A-31)corresponding to the plurality of interior helical splines 536 (seeFIGS. 30A, 31) of the press fit element 510. The at least onetranslational inerter assembly 500 (see FIGS. 29A-31) further comprisesthe torsion bar 514 (see FIGS. 29A-31) having the torsion bar first end564 a (see FIGS. 30A-30B) coupled to and installed in the inertiaelement 512 (see FIGS. 30A-30B), and having the torsion bar second end564 b (see FIGS. 30A-30B) fixedly attached to the support structure 116(see FIGS. 30B, 31) of the aircraft 100 (see FIGS. 25, 31). The torsionbar 514 (see FIGS. 30A-30B) further has the plurality of exterior linearsplines 572 (see FIGS. 30A-30B) corresponding to the plurality ofinterior linear splines 558 (see FIGS. 30A, 31) of the inertia element512 (see FIGS. 30A, 31).

The step 652 (see FIG. 32) of installing further comprises installingthe at least one translational inerter assembly 500 (see FIGS. 29A-31)further comprising a splined attachment element 520 (see FIGS. 29A-31)coupled to the torsion bar 514 (see FIGS. 29A-31). The torsion barsecond end 564 b (see FIGS. 30A-30B) is fixedly attached to the supportstructure 116 (see FIG. 30B) of the aircraft 100 (see FIGS. 25, 31), viathe first torsion bar bearing 506 a (see FIGS. 30B, 31) attached to thesplined attachment element 520 (see FIGS. 30B, 31). The first torsionbar bearing 506 a (see FIGS. 30B, 31) constrains the torsion bar 514(see FIGS. 30B, 31) in six (6) degrees of freedom. As shown in FIG. 29B,the first torsion bar 514 a is attached via the first torsion barbearing 506 a to the support structure 116, and the second torsion bar514 b is attached via the second torsion bar bearing 506 b to thesupport structure 116.

The step 652 (see FIG. 32) of installing further comprises installingthe at least one translational inerter assembly 500 (see FIGS. 29A-31)further comprising attachment points 504 (see FIGS. 29A-31), such as thefirst bearing 504 a (see FIGS. 29A-31) attaching the first end 502 a(see FIGS. 29A-31) of the flight control surface 122 (see FIGS. 29A-31)to the support structure 116 (see FIGS. 29A-29B, 30B, 31), and thesecond bearing 504 b (see FIGS. 29A-31) attaching the second end 502 b(see FIGS. 29A-31) of the flight control surface 122 to the supportstructure 116. The first bearing 504 a (see FIGS. 29A-31) and the secondbearing 504 b (see FIGS. 29A-31) constrain the flight control surface122 (see FIGS. 29A-31) in five (5) degrees of freedom and allow rotationabout the hinge axis 126 (see FIGS. 29A, 30A).

As shown in FIG. 32, the method 650 further comprises step 654 ofrotating the flight control surface 122 (see FIGS. 29A-31) using one ormore actuators 202 (see FIGS. 28, 31), such as hydraulic actuators 204(see FIG. 28). As shown in FIGS. 28 and 31, the actuator 202 maycomprise the cap end 212, the piston 216, the piston rod 224, and therod end 214. The rod end 214 (see FIG. 28) may be coupled to the onetranslational inerter assembly 500 (see FIGS. 27, 28) via the bellcrank128 (see FIG. 28) or via the horn element 516 (see FIGS. 27, 31) andbearing 518 (see FIGS. 27, 31).

As shown in FIG. 32, the method 650 further comprises step 656 using theat least one translational inerter assembly 500 (see FIGS. 29A-31) toaxially accelerate in a translational movement 604 (see FIG. 31) alongthe hinge axis 126 (see FIGS. 29A, 30A, 31) of the flight controlsurface 122 (see FIGS. 29A-31), the inertia element 512 (see FIGS.29A-31) relative to the torsion bar 514 (see FIGS. 29A-31), simultaneouswith, and in proportion to, rotation of the flight control surface 122.

The step 656 (see FIG. 32) of using the at least one translationalinerter assembly 500 (see FIGS. 29A-31) to axially accelerate theinertia element 512 (see FIGS. 29A-31) relative to the torsion bar 514(see FIGS. 29A-31) comprises axially accelerating the first inertiaelement 512 a (see FIGS. 29B, 31) and the second inertia element 512 b(see FIGS. 29B, 31) in opposite directions 553 (see FIG. 31), along therespective first torsion bar 514 a (see FIGS. 29B, 31) and the secondtorsion bar 514 b (see FIGS. 29B, 31), when the flight control surface122 (see FIGS. 29B, 31) is rotated.

As shown in FIG. 32, the method 650 further comprises step 658 ofdamping movement 606 (see FIG. 31) of the flight control surface 122(see FIGS. 29A-31), using the at least one translational inerterassembly 500 (see FIGS. 29A-31). The step 658 (see FIG. 32) of dampingmovement 606 (see FIG. 31) of the flight control surface 122 (see FIGS.29A-31) further comprises damping movement 606 of the flight controlsurface 122 to provide increased flutter suppression 610 (see FIG. 31)of the flight control surface 122. This preferably results in animproved hydraulic application stability 612 (see FIG. 31) and anincreased efficient flight control actuation 614 (see FIG. 31).

As shown in FIG. 32, the method 650 may further comprise step 660 of,after the step 652 of installing and prior to the step 654 of rotating,optionally, installing another translational inerter assembly 500 (seeFIGS. 29B, 31) at the second end 502 b (see FIGS. 29B, 31) of the flightcontrol surface 122 (see FIGS. 29B, 31). The two translational inerterassemblies 500 (see FIGS. 29B, 31) preferably comprise the firsttranslational inerter assembly 500 a (see FIGS. 29B, 31), with the firstpress fit element 510 a (see FIGS. 29B, 31), the first inertia element512 a (see FIGS. 29B, 31), and the first torsion bar 514 a (see FIGS.29B, 31), attached to the first end 502 a (see FIGS. 29B, 31) of theflight control surface 122, and preferably comprise the secondtranslational inerter assembly 500 b (see FIGS. 29B, 31), with thesecond press fit element 510 b (see FIGS. 29B, 31), the second inertiaelement 512 b (see FIGS. 29B, 31), and the second torsion bar 514 b (seeFIGS. 29B, 31), attached to the second end 502 b (see FIGS. 29B, 31) ofthe flight control surface 122.

Disclosed versions of the translational inerter assembly 500 (see FIGS.29A-31) and method 650 (see FIG. 32) for damping movement 606 (see FIG.31) of the flight control surface 122 (see FIGS. 29A-31) of the aircraft100 (see FIGS. 25, 31) have numerous advantages such as, increasedflutter suppression 610 (see FIG. 31) by providing surface torsioncritical damping, and reduced actuator swept volume and smalleractuators providing engine power offtake and weight and power offtakereduction, where reduced power offtake enables more fuel efficientengines. In addition, disclosed versions of the translational inerterassembly 500 (see FIGS. 29A-31) and method 650 (see FIG. 32) have theadvantages of enabling high pressure hydraulic actuator sizing to reduceflow and weight, reducing space required by enabling smaller actuator,valve, and horn element radius, reducing required stiffness to achievean aero-servo-elasticity goal, addressing force equalization by enablingreduced linear stiffness, increasing compliance that reduces force-fightfor active-active modes, enhancing electro-hydrostatic actuator (EHA)bandwidth without increasing stiffness or heat, and reducing flightcontrol surface resonant amplitude that reduces fatigue. Moreover,disclosed versions of the translational inerter assembly 500 (see FIGS.29A-31) and method 650 (see FIG. 32) have the further advantages ofenabling unanticipated kinematic action to further reduce flow andweight, reducing fatigue by eliminating amplification of torsionresonance, resolving potential electro-hydrostatic actuator (EHA)thermal issues by relaxing stiffness requirement, enabling relaxation ofan electromechanical actuator (EMA) backlash requirement, includingtolerances in automated sizing to evaluate required precision, usingvariable pressure to achieve weight reduction at comparable reliability,and sizing pressure, e.g., greater than 6000 psi (pounds per squareinch) which may decrease offtake with small weight penalty, where lowerofftake enables higher bypass ratio, which improves fuel efficiency.

In addition, disclosed versions of the translational inerter assembly500 (see FIGS. 29A-31) and method 650 (see FIG. 32) permit theelimination of any stiffness constraint in sizing, which enables reducedhydraulic system and aircraft size, reduced weight, and reduced power.This solution presents a more space efficient assembly and method.Further, disclosed versions of the translational inerter assembly 500(see FIGS. 29A-31) and method 650 (see FIG. 32) may improve theperformance of aircraft 100 (see FIG. 25) by improving the actuationsystem design. Further, the translational inerter assembly 500 (seeFIGS. 30A-30B) and method 650 (see FIG. 32) is a passive solution 616(see FIG. 31), that is, it uses a passive means to change the dynamicsof the flight control system 120 (see FIG. 25) instead of the activecontrol elements such as the actuator 202 (see FIG. 28) and valve sizeor diameter.

Moreover, disclosed versions of the translational inerter assembly 500(see FIGS. 29A-31) and method 650 (see FIG. 32) may be installed withinthe flight control surface 122 (see FIGS. 29A-31), about the hinge axis126 (see FIG. 29A, 30A) on the aircraft 100 (see FIGS. 25, 31), toincrease damping of the flight control surface 122. Increased dampingwill suppress flutter response of the flight control surface 122 (seeFIGS. 29A-31) to provide improved hydraulic application stability 612(see FIG. 31) and enable an increased efficient flight control actuation614 (see FIG. 31). This solution permits the addition of one or moretranslational inerter assemblies 500 (see FIGS. 29A-31) to the flightcontrol system 120 (see FIGS. 25, 31), which changes the dynamiccharacteristic of the hardware under control, rather than complicatingthe control elements themselves.

In addition, disclosed versions of the translational inerter assembly500 (see FIGS. 30A-30B) and method 650 (see FIG. 32) may be used toaddress flutter critical control surface applications on aircraft 100(see FIG. 25) to further optimize the translational inerter assembly 500(see FIGS. 30A-30B) design by enabling a translating inerter instead ofa rotational inerter, where the translational inerter assembly 500 (seeFIGS. 30A-30B) comprises the torsion bar 514 (see FIGS. 30A-30B), ortorsion rod, fixed to the support structure 116 (see FIGS. 26, 30B),such as the wing 114 (see FIGS. 26, 30B), the wing portion supportstructure 17 (see FIGS. 26, 30B), and/or the wing spar 18 (see FIG. 26),and where the inertia element 512 (see FIGS. 30A-30B) translates ratherthan rotates. The linear spline exterior 570 (see FIGS. 30A-30B) of thetorsion bar 514 (see FIGS. 30A-30B) is undercut at its fatigue locationand preferably has a longitudinal linear spline exterior 570 (see FIG.30A), where the inertia element 512 (see FIGS. 30A-30B) slidestranslationally along the plurality of exterior linear splines 572 (seeFIGS. 30A-30B), whenever the flight control surface 122 (see FIGS.30A-30B) rotates.

The inertia element 512 (see FIGS. 30A-30B) is accelerated only in thetranslational movement 604 (see FIG. 31), or fashion, instead ofrotating. Preferably, there are also at least two attachment points 504(see FIGS. 30A-30B) connected to the support structure 116 (see FIGS.26, 30B), such as the wing 114 (see FIGS. 26, 30B), the wing portionsupport structure 17 (see FIGS. 26, 30B), and/or the wing spar 18 (seeFIG. 26), for further structural robustness. The torsion bar 514 (seeFIGS. 29A-31) is extended beyond the first end 502 a (see FIG. 29A)and/or second end 502 b (see FIG. 29B) of the flight control surface 122(see FIGS. 29A-31) and supported by the support structure 116 (see FIGS.29A-29B, 30B, 31) in all six (6) degrees of freedom. Thus, the rotationof the flight control surface 122 (see FIGS. 29A-31) causes each inertiaelement 512 (see FIGS. 29A-31), for example, the first inertia element512 a (see FIG. 29B) and the second inertia element 512 b (see FIG.29B), to translate along the linear spline exterior 570 (see FIG. 30A)of the torsion bar 514 (see FIG. 30A) in opposite directions 553 (seeFIG. 31). This opposite motion is required to reduce reaction force onthe flight control surface 122 (see FIGS. 29A-31) from loading wingstructure. When the flight control surface 122 (see FIGS. 29A-31) andthe fixed press fit element 510 (see FIGS. 29A-31) within the flightcontrol surface 122 moves, the inertia element 512 (see FIGS. 29A-31) isforced to translate along the hinge axis 126 (see FIGS. 29A, 30A, 31) bythe helical or threaded engagement of the flight control surface 122(see FIGS. 29A-31) and the fixed press fit element 510 (see FIGS.29A-31) and the inertia element 512. This acceleration will dampenflight control surface 122 motion. Optionally, a second torsion bar 514b (see FIG. 29B) may be added on the other side of the flight controlsurface 122 to improve surface dampening. In addition, the first torsionbar bearing 506 a (see FIGS. 29A-29B) constrains the torsion bar 514,such as the first torsion bar 514 a (see FIGS. 29A-29B), and the secondtorsion bar bearing 506 b (see FIG. 29B) constrains the torsion bar 514,such as the second torsion bar 514 b (see FIG. 29B), in six (6) degreesof freedom. The first bearing 504 a (see FIGS. 29A-2B) and the secondbearing 504 b (see FIG. 29B) constrain surface motion of the flightcontrol surface 122 (see FIGS. 29A-29B) in five (5) degrees of freedomto permit rotation about the hinge axis 126 (see FIGS. 29A, 30A, 31).

Additional modifications and improvements of the present disclosure maybe apparent to those of ordinary skill in the art. Thus, the particularcombination of parts described and illustrated herein is intended torepresent only certain examples of the present disclosure and is notintended to serve as limitations of alternative examples or deviceswithin the spirit and scope of the disclosure.

What is claimed is:
 1. A translational inerter assembly for dampingmovement of a flight control surface of an aircraft with a supportstructure, the translational inerter assembly comprising: a press fitelement rotatably disposed within the flight control surface; an inertiaelement disposed in the press fit element; and a torsion bar coupled tothe inertia element and to the support structure of the aircraft, suchthat when the flight control surface rotates, the inertia elementtranslates, and movement of the flight control surface is dampened. 2.The translational inerter assembly of claim 1, further comprising asplined attachment element coupled to the torsion bar, the splinedattachment element having a plurality of interior linear splinescorresponding to a plurality of exterior linear splines of the torsionbar.
 3. The translational inerter assembly of claim 1, wherein the pressfit element comprises a plurality of interior helical splines, theinertia element comprises a plurality of exterior helical splines and aplurality of interior linear splines, and the torsion bar comprises aplurality of exterior linear splines, and further wherein the pluralityof exterior helical splines of the inertia element correspond to theplurality of interior helical splines of the press fit element, and theplurality of exterior linear splines of the torsion bar correspond tothe plurality of interior linear splines of the inertia element.
 4. Thetranslational inerter assembly of claim 1, wherein the translationalinerter assembly is attached to a horn and a bearing, and the horn andthe bearing are further attached to one or more actuators for surfaceposition control.
 5. The translational inerter assembly of claim 4,wherein the one or more actuators comprise one or more hydraulicactuators, each of the one or more actuators comprising a piston coupledto a piston rod, the piston slidable within an actuator housing, andfurther comprising a rod end and a cap end axially movable relative toone another.
 6. The translational inerter assembly of claim 1, whereinthe torsion bar is fixedly attached to the support structure of theaircraft, via a first torsion bar bearing, the first torsion bar bearingconstraining the torsion bar in six (6) degrees of freedom.
 7. Thetranslational inerter assembly of claim 1, further comprising a firstbearing attaching a first end of the flight control surface to thesupport structure, and a second bearing attaching a second end of theflight control surface to the support structure, wherein the firstbearing and the second bearing constrain the flight control surface infive (5) degrees of freedom and allow rotation about a hinge axis of theflight control surface.
 8. The translational inerter assembly of claim1, wherein when the movement of the flight control surface is dampened,an increased flutter suppression of the flight control surface isobtained, resulting in an improved hydraulic application stability andan increased efficient flight control actuation.
 9. The translationalinerter assembly of claim 1, wherein the flight control surfacecomprises an aileron, and the support structure comprises one of, awing, a wing portion support structure, and a wing spar.
 10. Anaircraft, comprising: a flight control surface pivotably coupled to asupport structure; one or more actuators actuating the flight controlsurface; and at least one translational inerter assembly for dampingmovement of the flight control surface, the at least one translationalinerter assembly comprising: a press fit element rotatably disposedwithin the flight control surface; an inertia element disposed in thepress fit element; and a torsion bar coupled to the inertia element andto the support structure, such that when the flight control surfacerotates, the inertia element translates, and movement of the flightcontrol surface is dampened.
 11. The aircraft of claim 10, wherein theaircraft comprises two translational inerter assemblies comprising afirst translational inerter assembly attached to a first end of theflight control surface, and comprising a second translational inerterassembly attached to a second end of the flight control surface.
 12. Theaircraft of claim 11, wherein the first translational inerter assemblyhas a first inertia element with a plurality of exterior helical splinescomprising right hand direction helical splines, and wherein the secondtranslational inerter assembly has a second inertia element with aplurality of exterior helical splines comprising left hand directionhelical splines, and further wherein rotation of the flight controlsurface causes the first inertia element and the second inertia elementto translate along a hinge axis of the flight control surface, and alonga linear spline exterior of the torsion bar in opposite directions, toreduce reaction force on the flight control surface.
 13. The aircraft ofclaim 10, wherein the flight control surface comprises an aileron, andthe support structure comprises one of, a wing, a wing portion supportstructure, and a wing spar.
 14. The aircraft of claim 10, wherein thetranslational inerter assembly further comprises a first bearingattaching a first end of the flight control surface to the supportstructure, and a second bearing attaching a second end of the flightcontrol surface to the support structure, and further wherein the firstbearing and the second bearing constrain the flight control surface infive (5) degrees of freedom and allow rotation about a hinge axis of theflight control surface.
 15. A method for damping movement of a flightcontrol surface of an aircraft having a support structure, the methodcomprising the steps of: attaching at least one translational inerterassembly to the flight control surface, the at least one translationalinerter assembly comprising: a press fit element rotatably disposedwithin the flight control surface; an inertia element disposed in thepress fit element; and a torsion bar coupled to the inertia element andto the support structure of the aircraft; rotating the flight controlsurface using one or more actuators, to cause the inertia element, viathe press fit element, to translate along a hinge axis of the flightcontrol surface, and to translate along the torsion bar; and dampingmovement of the flight control surface, using the at least onetranslational inerter assembly.
 16. The method of claim 15, furthercomprising, after attaching the at least one translational inerterassembly and prior to rotating the flight control surface, attachinganother translational inerter assembly to the flight control surface,wherein the two translational inerter assemblies comprise a firsttranslational inerter assembly, with a first press fit element, a firstinertia element, and a first torsion bar, attached to a first end of theflight control surface, and comprise a second translational inerterassembly, with a second press fit element, a second inertia element, anda second torsion bar, attached to a second end of the flight controlsurface.
 17. The method of claim 16, wherein rotating the flight controlsurface further comprises using the at least one translational inerterassembly to axially accelerate the first inertia element and the secondinertia element in opposite directions, along the respective firsttorsion bar and the second torsion bar, when the flight control surfaceis rotated.
 18. The method of claim 15, wherein rotating the flightcontrol surface using one or more actuators, to cause the inertiaelement to translate along the torsion bar further comprises rotatingthe flight control surface using one or more actuators, to cause theinertia element to translate along a linear spline exterior of thetorsion bar, the linear spline exterior having a plurality of exteriorlinear splines corresponding to a plurality of interior linear splinesof the inertia element.
 19. The method of claim 15, wherein rotating theflight control surface using one or more actuators, comprises using oneor more hydraulic actuators each comprising a piston coupled to a pistonrod, the piston slidable within an actuator housing, and furthercomprising a rod end and a cap end axially movable relative to oneanother, and further wherein the one or more actuators are attached to ahorn and a bearing, which is attached to the at least one translationalinerter assembly.
 20. The method of claim 15, wherein damping movementof the flight control surface further comprises damping movement of theflight control surface to provide increased flutter suppression of theflight control surface, resulting in an improved hydraulic applicationstability and an increased efficient flight control actuation.